Composite component having additively printed inner portion

Abstract

To provide a composite component having an additively printed inner portion.SOLUTION: A composite component 100 for a gas turbine engine 10 is provided, along with methods of forming the same. The composite component 100 includes an additively printed inner portion 120 defining at least one flowpath feature 124, and a ceramic matrix composite (CMC) outer portion 122. The CMC outer portion 122 is formed on the additively printed inner portion 120 such that the CMC outer portion 122 substantially surrounds the additively printed inner portion 120. The additively printed inner portion 120 includes SiC, and the CMC outer portion 122 includes a fiber-reinforced ceramic matrix (e.g., including SiC) and defines at least one cooling cavity fluidly coupled to the at least one flowpath feature 124 of the additively printed inner portion 120.SELECTED DRAWING: Figure 2A

Description

[Technical field]

[0001] The present disclosure relates to a composite component that includes an additively printed interior portion. [Background technology]

[0002] Reinforced ceramic matrix composites ("CMCs"), which include fibers dispersed in a continuous ceramic matrix of the same or different composition, are well suited for structural applications due to their toughness, heat resistance, high temperature strength, and chemical stability. Such composites typically have a high strength-to-weight ratio, which makes them attractive in applications where weight is important (e.g., aeronautical applications, etc.). Their stability at high temperatures makes CMCs very suitable for applications where the components are in contact with hot gases (e.g., gas turbine engines, etc.). However, the ceramic fibers used to form fiber-reinforced CMC materials can be expensive. Additionally, it can be difficult to cool fiber-reinforced composites. Improved components that address one or more of these challenges would therefore be desirable. Summary of the Invention [Means for solving the problem]

[0003] A full and enabling disclosure of this disclosure, including the best mode thereof, directed to one of ordinary skill in the art is set forth in this specification, which makes reference to the accompanying figures. [Brief description of the drawings]

[0004] [Figure 1] 1 is a schematic cross-sectional view of a gas turbine engine according to an exemplary aspect of the present disclosure; [Figure 2A] FIG. 2 is a perspective view of a composite airfoil having an additively printed inner portion according to an exemplary embodiment of the present disclosure; [Figure 2B] FIG. 2B is a perspective view of the additionally printed inner portion of FIG. 2A. [Diagram 3] FIG. 2 is a schematic cross-sectional view of a composite airfoil having an additively printed inner portion positioned at a trailing edge region of the composite airfoil according to an exemplary aspect of the present disclosure; [Figure 4] FIG. 2 is a schematic cross-sectional view of a composite airfoil having an additive printed inner portion defining an additive printed inner portion cavity, according to an exemplary embodiment of the present disclosure; [Diagram 5] FIG. 2 is a schematic cross-sectional view of a composite airfoil having an additively printed inner portion positioned at or near a leading edge of the composite airfoil, according to an exemplary embodiment of the present disclosure; [Figure 6] FIG. 2 is a schematic cross-sectional view of a composite airfoil having an additive printed inner portion defining an additive printed inner portion manifold that is divided into multiple chambers, according to an exemplary embodiment of the present disclosure. [Figure 7] FIG. 2 is a schematic cross-sectional view of a composite airfoil having an additively printed inner portion disposed along a suction side of the composite airfoil in accordance with an exemplary embodiment of the present disclosure; [Figure 8A] FIG. 2 is a schematic cross-sectional view of a composite airfoil having an additively printed inner portion defining a serpentine cavity, according to an exemplary embodiment of the present disclosure; [Figure 8B] FIG. 8C is a schematic cross-sectional view of the composite airfoil of FIG. 8B, the cross-section being taken along a radial direction. [Figure 9] FIG. 2 is a schematic cross-sectional view of a composite nozzle airfoil assembly having an additively printed inner portion disposed within an inner band according to an exemplary aspect of the present disclosure; [Figure 10] 1 is a schematic cross-sectional view of a composite shroud having an additively printed inner portion in accordance with an exemplary embodiment of the present disclosure; FIG. [Figure 11A] 1 is a schematic cross-sectional view of a composite airfoil having an additively printed inner portion defining a cellular structure, according to an exemplary embodiment of the present disclosure; FIG. [Figure 11B]FIG. 11B is a schematic cross-sectional view of the composite airfoil of FIG. 11A, the cross-section being taken along a radial direction. [Figure 12] FIG. 1 is a schematic cross-sectional view of a composite shroud having an additively printed inner portion that served as a mandrel, according to an exemplary embodiment of the present disclosure. [Figure 13] 1 is a flow diagram illustrating a method for forming a composite material component according to an exemplary embodiment of the present disclosure. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0005] definition The word "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. Additionally, unless specifically identified otherwise, all embodiments described herein should be considered as exemplary.

[0006] The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise.

[0007] For example, the term "at least one of" in the context of "at least one of A, B, and C" refers to A only, B only, C only, or any combination of A, B, and C.

[0008] The term "turbomachine" or "turbomachine" refers to a machine that includes one or more compressors, a heat generation section (e.g., a combustion section), and one or more turbines that together produce a torque output.

[0009] The term "gas turbine engine" refers to an engine that has a turbomachine as all or part of its power source. Exemplary gas turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, and the like, as well as hybrid-electric versions of one or more of these engines.

[0010] The terms "upstream" and "downstream" refer to relative directions with respect to fluid flow in a fluid pathway. For example, "upstream" refers to the direction from which fluid flows, and "downstream" refers to the direction to which fluid flows.

[0011] As used herein, the terms "axial" and "axially" refer to directions and orientations that extend substantially parallel to the centerline of a gas turbine engine. Additionally, the terms "radial" and "radially" refer to directions and orientations that extend substantially perpendicular to the centerline of a gas turbine engine. Additionally, as used herein, the terms "circumferential" and "circumferentially" refer to directions and orientations that extend arcuately about the centerline of a gas turbine engine.

[0012] Terms such as "coupled," "fixed," and "attached" refer to both a direct connection, fastening, or attachment, and an indirect connection, fastening, or attachment through one or more intermediate components or features, unless otherwise specified in this specification.

[0013] As used herein, the terms "first," "second," and "third" may be used interchangeably to distinguish one component from another and are not intended to denote the location or importance of the individual components.

[0014] For purposes of the following description, the terms "upper", "lower", "right", "left", "vertical", "horizontal", "top", "bottom", "lateral", "longitudinal", and their derivatives shall refer to the embodiments as they are oriented in the drawings. However, it should be understood that the embodiments are capable of various alternative variations unless expressly specified to the contrary. It should also be understood that the specific devices illustrated in the accompanying drawings (and described in the following specification) are merely exemplary embodiments of the present disclosure. Hence, specific dimensions and other physical attributes related to the embodiments disclosed herein are not to be considered as limiting.

[0015] The term "adjacent" as used herein in reference to two walls and / or surfaces refers to the two walls and / or surfaces being in contact with one another, or the two walls and / or surfaces being separated only by one or more nonstructural layers, as well as the two walls and / or surfaces and the one or more nonstructural layers being in series contact (i.e., a first wall / surface in contact with one or more nonstructural layers and one or more nonstructural layers in contact with a second wall / surface).

[0016] As used herein, the terms "integral," "unitary," or "monolithic" as used to describe a structure refer to the structure being integrally formed from a continuous material or group of materials without seams or connecting joints, etc. The integral unitary structures described herein can be formed through additive manufacturing to have the described structure, or alternatively, through ply layup processes, casting processes, etc.

[0017] As used herein, the term "build plane" refers to the plane defined by the surface upon which an energy beam impinges to selectively irradiate and thereby consolidate powder material during an additive manufacturing process.

[0018] As used herein, the terms "consolidate" or "consolidating" refer to the solidification of a powder material (including by melting, dissolving, or sintering, etc.) as a result of irradiating the powder material.

[0019] As used herein, ceramic matrix composites or "CMC" refers to a class of materials that include 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 matrix materials for CMCs include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), oxide ceramics (e.g., silicon oxycarbide, silicon oxynitride, aluminum oxide (Al 2 O 3 ), silicon dioxide (SiO 2 ), aluminosilicates, or mixtures thereof. 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 be included in the CMC matrix.

[0020] Some examples of reinforcing fibers for CMCs include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), non-oxide carbon-based materials (e.g., carbon), oxide ceramics (e.g., silicon oxycarbide, silicon oxynitride, aluminum oxide (Al 2 O 3 ), silicon dioxide (SiO 2), aluminosilicates such as mullite, or mixtures thereof), or mixtures thereof.

[0021] As used herein, the terms "substantially surround" and "substantially encircling" mean that the inner portion has at least 90% of its surface area covered by the CMC outer portion (e.g., 95% of its surface area covered by the CMC outer portion). It is noted that the area defined by any inlet portion 126 and / or outlet portion 128 is not included in the surface area of ​​the inner portion 120. The terms "substantially surround" and "substantially encircling" may include completely surrounding (e.g., encircling) and completely surrounding (e.g., encircling), respectively. The terms "surround" and "encircling" refer to completely surrounding the inner portion except for any fluid cavities or channels defined therein.

[0022] Detailed Description of the Preferred Embodiments Reference will now be made in detail to the present embodiments of the disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and the description are used to refer to like or similar parts of the disclosure.

[0023] In general, certain CMCs can be referred to as their fiber type / matrix type combinations, e.g., 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 silicon carbide fiber reinforced silicon carbide / silicon nitride matrix mixture, etc. In other examples, CMCs can be oxide-based materials (e.g., aluminum oxide (Al 2 O 3 ), silicon dioxide (SiO 2The aluminosilicates can be made up of a matrix and reinforcing fibers including crystalline materials (e.g., mullite (3Al), aluminosilicates, and mixtures thereof). 2 O 3 2SiO 2 ) and glassy aluminosilicates.

[0024] In certain embodiments, the reinforcing fibers can be bundled and / or coated before being incorporated into the matrix. For example, the fiber bundles can be formed as a reinforcing tape (e.g., unidirectional reinforcing tape, etc.). Multiple tapes can be laid up together to form a preform component. The fiber bundles can be impregnated with a slurry composition before or after forming the preform. The preform can then be subjected to a heat treatment (e.g., curing or burnout, etc.) to produce high carbide residue in the preform and a subsequent chemical treatment (e.g., melt infiltration with silicon, etc.) to arrive at a component formed from a CMC material having a desired chemical composition.

[0025] Such materials, along with certain monolithic ceramics (i.e., ceramic materials without reinforcing materials), are particularly suitable for higher temperature applications. Additionally, these ceramic materials are lighter weight compared to superalloys, yet can provide strength and durability to components made therefrom. Thus, such materials are currently being considered for many gas turbine components used in the higher temperature sections of gas turbine engines, such as, for example, airfoils (e.g., turbines, and vanes), combustors, shrouds, and other similar components, which would benefit from the lighter weight and higher temperature capabilities that these materials can provide.

[0026] As described herein, the presently disclosed subject matter involves the use of an additive manufacturing machine or system. As used herein, the term "additive manufacturing" generally refers to a manufacturing technique in which components are manufactured in a layer-by-layer manner. An exemplary additive manufacturing machine can be configured to utilize any suitable additive manufacturing technique. The additive manufacturing machine can utilize additive manufacturing techniques including powder bed fusion (PBF) techniques (e.g., direct metal laser melting (DMLM) techniques, selective laser melting (SLM) techniques, direct metal laser sintering (DMLS) techniques, or selective laser sintering (SLS) techniques, etc.). In an exemplary PBF technique, thin layers of powder material are applied sequentially to a build plane and then selectively melted or fused to one another in a layer-by-layer manner to form one or more three-dimensional objects. Additively manufactured objects are generally monolithic in nature and can have various integral subcomponents.

[0027] Additionally or alternatively, suitable additive manufacturing techniques may include, for example, Fused Deposition Modeling (FDM) techniques, Direct Energy Deposition (DED) techniques, Laser Engineered Net Shaping (LENS) techniques, Laser Net Shape Manufacturing (LNSM) techniques, Direct Metal Deposition (DMD) techniques, Digital Light Processing (DLP) techniques, binder jet printing, and other additive manufacturing techniques that utilize an energy beam or other energy source to solidify additive manufacturing materials, such as powder materials, etc. Indeed, any suitable additive manufacturing modality may be utilized with the presently disclosed subject matter.

[0028] Additive manufacturing techniques can generally be described as the fabrication of an object by building the object up point-by-point, line-by-line, layer-by-layer, typically in a vertical direction. Other fabrication methods are contemplated and are within the scope of the present disclosure. For example, although the discussion herein refers to the addition of material to form successive layers, the presently disclosed subject matter can be practiced by any additive or other manufacturing technique, including layer-additive, layer-subtractive, or hybrid processes.

[0029] The additive manufacturing processes described herein can be used to form components using any suitable material. For example, the material can be a metal, ceramic, polymer, epoxy, photopolymer resin, plastic, or any other suitable material that can be a solid, powder, sheet material, wire, or any other suitable form, or combinations thereof. Additionally or in the alternative, exemplary materials may include metals, ceramics, or binders, and combinations thereof. Exemplary ceramics may include ultra-high temperature ceramics and / or precursors for ultra-high temperature ceramics (e.g., polymeric precursors, etc.). Each successive layer can be, for example, between about 10 μm and 200 μm, although the thickness can be determined based on any number of parameters and can be any suitable size.

[0030] Generally, the surface of the powder bed defines a build plane. During irradiation of each layer of the powder bed, a previously irradiated portion of each layer may define a portion of the build plane. Prior to dispensing powder material across the build module, a build plate that supports the powder bed generally defines the build plane.

[0031] The present disclosure generally relates to additively printed interior portions for composite material components to define one or more passages (e.g., cooling channels, etc.) therein. For example, an additively printed interior portion for a composite material component can define one or more passages, and then the additively printed interior portion can be rolled or disposed within a composite material to form the composite material component. The additively printed interior portion (with one or more passages defined therein) can be formed from additive manufacturing processes, etc., to define complex features or geometries that are difficult or impossible to form in the composite material component using conventional techniques (e.g., laser machining, electrical discharge machining (EDM), etc.). As one example, the present subject matter enables cooling features for improved cooling of airfoils (e.g., improved cooling of the airfoil trailing edge region by introducing passages closer to the trailing edge of the airfoil and by increasing the heat transfer coefficient over the cold side or suction side while avoiding the performance penalty that results from film cooling the suction side). Thus, the additively printed inner portions described herein can improve cooling of a composite component, for example, by defining passageways for cooling fluid that provide targeted cooling of particular features or locations within or on the composite component, and / or by increasing the heat transfer coefficient or otherwise enabling more efficient heat transfer in one or more locations of the composite component.

[0032] Referring now to the drawings, where identical numerals denote the same elements throughout the figures, FIG. 1 is a schematic cross-sectional view of a gas turbine engine according to an exemplary aspect of the present disclosure. More specifically, with respect to the embodiment of FIG. 1, the gas turbine engine is a high bypass turbofan jet engine 10 (referred to herein as "turbofan engine 10"). As shown in FIG. 1, the turbofan engine 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 provided for reference), a circumferential direction C (extending about the longitudinal centerline 12 and the axial direction A), and a radial direction R. Generally, the turbofan engine 10 includes a fan section 14 and a core turbine engine 16 disposed downstream from the fan section 14.

[0033] The depicted exemplary core turbine engine 16 generally includes a substantially tubular outer casing 18 defining an annular inlet section 20. The outer casing 18 encloses, in serial-flow relationship, a compressor section including a booster 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 a jet exhaust nozzle section 32. A high pressure (HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HP compressor 24. A low pressure (LP) shaft or spool 36 drivingly connects the LP turbine 30 to the LP compressor 22.

[0034] For the embodiment depicted, the fan section 14 includes a fan 38 having a plurality of fan blades 40 coupled in a spaced apart fashion to a hub or disk 42. As depicted, the fan blades 40 extend generally outwardly along a radial direction R from the disk 42. The fan blades 40 and disk 42 are rotatable together about the longitudinal centerline 12 by the LP shaft or spool 36. In some embodiments, a power gearbox having multiple gears can be included to reduce the rotational speed of the LP shaft or spool 36 to a more efficient rotational fan speed.

[0035] Still referring to the exemplary embodiment of FIG. 1 , the disk 42 is covered by a rotatable front nacelle 48 that is aerodynamically contoured to drive airflow through the plurality of fan blades 40. Additionally, the exemplary fan section 14 includes an annular fan casing or outer nacelle 50 that circumferentially surrounds at least a portion of the fan 38 and / or core turbine engine 16. It should be appreciated that the fan case (nacelle) 50 can be configured to be supported relative to the core turbine engine 16 by a plurality of circumferentially spaced outlet guide vanes 52. Moreover, a downstream section 54 of the fan case 50 can extend over an outer portion of the core turbine engine 16 so as to define a bypass airflow passage 56 therebetween.

[0036] During operation of the turbofan engine 10, a volume of air 58 enters the turbofan engine 10 through an associated inlet 60 of the fan case 50 and / or fan section 14. As the volume of air 58 passes across the fan blades 40, a first portion of the air 58, as indicated by arrow 62, is directed or routed into the bypass air flow passage 56 and a second portion of the air 58, as indicated by arrow 64, is directed or routed into the LP compressor 22. The ratio between the first portion of air 62 and the second portion of air 64 is commonly known as the bypass ratio. The pressure of the second portion of air 64 is then increased as it is routed through the compressor section and into the combustion section 26, where it is mixed with fuel and combusted to provide combustion gases 66. More specifically, the compressor section includes the LP compressor 22 and the HP compressor 24, which may each include a plurality of compressor stages 80, with each stage 80 including both an annular array or circumferential row of stationary compressor vanes 82 (also referred to as compressor stator vanes 82) and an annular array or circumferential row of compressor rotating blades 84 (also referred to as compressor rotor blades 84) positioned immediately downstream of the compressor stator vanes 82. The plurality of compressor rotor blades 84 in the LP compressor 22 are coupled to the LP shaft or spool 36, and the plurality of compressor blades in the HP compressor 24 are coupled to the HP shaft or spool 34. A plurality of compressor stator vanes 82 in the LP compressor 22 are coupled to the compressor casing, a plurality of compressor stator vanes 82 in the HP compressor 24 are coupled to the compressor casing, and at least a portion of the HP compressor stator vanes 82 are coupled to a compressor casing 90. In some embodiments, the compressor casing 90 may extend through both the LP compressor 22 and the HP compressor 24 and support all of the compressor stator vanes 82.In other embodiments, the compressor casing 90 supports only a portion of the compressor stator vanes 82 and may support only a portion of the compressor stator vanes 82 in the HP compressor 24. As previously described, as the second portion 64 of the air passes through sequential stages of the compressor stator vanes 82 and compressor rotor blades 84, a volume of the air 64 is pressurized, i.e., the pressure of the air 64 is increased, prior to being combusted with fuel in the combustion section 26 to form combustion gases 66.

[0037] The combustion gases 66 are routed through the HP turbine 28 where a portion of the thermal and / or kinetic energy from the combustion gases 66 is extracted via sequential stages of HP turbine stator vanes 68 coupled to the outer casing 18 and HP turbine rotor blades 70 coupled to the HP shaft or spool 34, thus causing the HP shaft or spool 34 to rotate, thereby supporting operation of the HP compressor 24. The combustion gases 66 are then routed through the LP turbine 30 where a second portion of the thermal and kinetic energy is extracted from the combustion gases 66 via sequential stages of LP turbine stator vanes 72 coupled to the outer casing 18 and HP turbine rotor blades 74 coupled to the LP shaft or spool 36, thus causing the LP shaft or spool 36 to rotate, thereby supporting operation of the LP compressor 22 and / or rotation of the fan 38.

[0038] The combustion gases 66 are then routed through the jet exhaust nozzle section 32 of the core turbine engine 16 to provide propulsive thrust. Concurrently, the pressure of the first portion of the air 62 is substantially increased as the first portion of the air 62 is routed through the bypass air flow passage 56 before it is exhausted from a fan nozzle exhaust section 76 of the turbofan engine 10, also providing propulsive thrust. The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section 32 at least partially define a hot gas flow path 78 for routing the combustion gases 66 through the core turbine engine 16.

[0039] Although the gas turbine engine of FIG. 1 is depicted in a turboshaft configuration, it will be appreciated that the teachings of the present disclosure may be applied to other types of turbine engines, or more generally, turbomachines, and other shaft systems. For example, the turbine engine may be another suitable type of gas turbine engine, such as a turboprop, turbojet, turbofan, aeroderivative, etc. The present disclosure may also be applied to other types of turbomachines, such as steam turbine engines, etc. Additionally, the present disclosure may be applied to other types of composite components (e.g., those used in applications other than turbomachinery).

[0040] Figure 2A provides a perspective view of a composite component 100, shown as an airfoil, for a gas turbine engine, such as the turbofan engine 10 illustrated in Figure 1. That is, for the embodiments described herein, the composite component 100 is formed from a composite material (e.g., a ceramic matrix composite (CMC)).

[0041] 2A, the airfoil 101 includes a concave pressure side 102 opposite a convex suction side 104. The pressure side 102 is opposite the suction side 104, and the opposite pressure side 102 and suction side 104 of the airfoil 101 extend along a radial direction R over a span S from an inner end 106 to an outer end 108. The inner end 106 may be, for example, a root of a blade airfoil, and the outer end 108 may be, for example, a tip of the blade airfoil. The pressure side 102 and suction side 104 of the airfoil 101 extend along an axis A between a leading edge 110 and a trailing edge 112, i.e., the trailing edge 112 is opposite the leading edge 110 along the axis A. A leading edge 110 defines a forward end of the airfoil 101 and a trailing edge 112 defines an aft end of the airfoil 101 .

[0042] As shown in Figure 2A, the composite material component 100 includes an additively printed inner portion 120 and a CMC outer portion 122 that substantially surrounds the additively printed inner portion 120. Figure 2B provides a perspective view of the additively printed inner portion 120 shown in Figure 2A. With reference to Figure 2B, the additively printed inner portion 120 defines a plurality of cooling channels 124 (which serve as a plurality of flow path features 124), each of the plurality of cooling channels 124 having an inlet portion 126 and an outlet portion 128. The additively printed inner portion 120 also has an outer surface 130.

[0043] 2A , the CMC outer portion 122 substantially surrounds the additively printed inner portion 120 such that the CMC outer portion 122 is in contact with the outer surface 130. Additionally, the CMC outer portion 122 has a cover outer surface 132. In the embodiment shown in FIG. 2A , the outlet portion 128 of each cooling channel 124 extends to the cover outer surface 132, which defines the outer surface of the airfoil 101. That is, the outlet portion 128 of each cooling channel 124 is defined along the outer surface of the airfoil 101. However, in some embodiments, the outlet portion 128 of one or more of the cooling channels 124 can be defined in the cover outer surface 132 such that at least one (but not necessarily all) of the outlet portions 128 are defined in the cover outer surface 132.

[0044] As shown in FIG. 2A , in some embodiments, the airfoil 101 defines a cooling cavity 134 for receipt of a cooling fluid F, with the trailing edge 112 of the airfoil 101 defined by a trailing edge region 136 positioned aft of the cooling cavity 134. The cooling cavity 134 may be defined by the CMC outer portion 122 or another portion of the composite material. Furthermore, the cooling cavity 134 may extend along the radial direction R for all or only a portion of the span S between the inboard end 106 and the outboard end 108. Moreover, although the cooling cavity 134 is shown in FIG. 2A (and other figures) as being defined adjacent the leading edge 110, it will be appreciated that the cooling cavity 134 may be defined at any suitable location along the axial direction A between the leading edge 110 and the trailing edge 112.

[0045] 2A , the additive printed inner portion 120 is disposed in the trailing edge region 136 and positioned adjacent the suction side 104 such that the additive printed inner portion 120 is closer to the suction side 104 than to the pressure side 102. More generally, a midline M ( FIG. 2B ) of the additive printed inner portion 120 extending along the axial direction A can be closer to one of the pressure side 102 or the suction side 104 than to the other of the pressure side 102 or the suction side 104.

[0046] Positioning the additive printed inner portion 120 closer to one side of the airfoil 101 than the other side of the airfoil 101 can allow for tailored or targeted cooling of the airfoil 101. For example, as shown in FIG. 2A , by disposing the additive printed inner portion 120 adjacent to the suction side 104, the suction side 104 can be cooled through heat transfer with cooling fluid F flowing through cooling channels 124 of the additive printed inner portion 120, which can eliminate the need to film cool the suction side, thereby eliminating the performance penalty associated with such film cooling. Additionally or alternatively, in some embodiments, the additive printed inner portion 120 can be positioned closer to the pressure side 102 than the suction side 104 to provide the needed cooling axially along the pressure side 102.

[0047] In some embodiments (such as that shown in FIG. 2A ), the inlet 126 of each cooling channel 124 can be fluidly coupled to a cooling cavity 134 of the component 100, such that, for example, a cooling fluid F can flow from the cooling cavity 134 into the cooling channel 124, and the cooling fluid F exits each cooling channel 124 at an outlet 128 of the respective cooling channel 124. Referring to FIG. 3 , in other embodiments, the additively printed inner portion 120 can be spaced apart from the cooling cavity 134, such that the inlet 126 of each cooling channel 124 is defined apart from the cooling cavity 134 while remaining fluidly coupled together. In such an embodiment, a feed channel 138 extends from the cooling cavity 134 to the inlet 126 of each cooling channel 124 of the plurality of cooling channels 124. As such, for the additively printed inner portion 120 defining a plurality of cooling channels 124, the airfoil 101 may define a corresponding plurality of feed channels 138 within each one of the feed channels 138 extending between the cooling cavity 134 and each one of the cooling channels 124 of the additively printed inner portion 120.

[0048] Each feed channel 138 includes a feed channel inlet 140 and a feed channel outlet 142. The feed channel inlet 140 is defined in the cooling cavity 134, and the feed channel outlet 142 is defined in a respective cooling channel 124. For example, the feed channel outlet 142 of each feed channel 138 can extend to the inlet 126 of a respective cooling channel 124. In this manner, each feed channel 138 defines a passage or pathway for the cooling fluid F to flow from the cooling cavity 134 to the additively printed inner portion 120 (e.g., for the embodiment of FIG. 3, to the cooling channel 124 of the additively printed inner portion 120). In the embodiment of FIG. 3, the inner portion 120 is enclosed within the CMC outer portion 122.

[0049] The one or more feed channels 138 may be defined using conventional techniques. For example, the one or more feed channels 138 may be laser drilled, may be defined by electrical discharge machining (EDM), may be formed through the use of a fugitive material (which is burned out or otherwise removed during thermal and / or chemical treatment of the composite), or may be formed by any other suitable technique or combination of techniques.

[0050] 4, another embodiment of an additive printed inner portion 120 for a composite material component such as an airfoil 101 is described, where the inner portion 120 is enclosed within a CMC outer portion 122. As shown in the schematic cross-sectional view of FIG. 4, the additive printed inner portion 120 may be positioned aft of the cooling cavity 134, and the additive printed inner portion 120 may define an additive printed inner portion cavity 144 extending along at least a portion of the span S (FIG. 2A). That is, the additive printed inner portion 120 may define its own cooling cavity (additive printed inner portion cavity 144) for receipt of a cooling fluid F extending along the radial direction R (FIG. 2A) across at least a portion of the span S between the inner end 106 and the outer end 108 of the airfoil 101.

[0051] As further illustrated in Figure 4, at least one feed channel 138 extends from the cooling cavity 134 to the additively printed inner portion cavity 144. For example, for each feed channel 138, a feed channel inlet portion 140 is defined in the cooling cavity 134 and a feed channel outlet portion 142 is defined in the additively printed inner portion cavity 144. As such, each feed channel 138 defines a passage or pathway for cooling fluid F to flow from the cooling cavity 134 to the additively printed inner portion 120 (e.g., for the embodiment of Figure 4, to the additively printed inner portion cavity 144 of the additively printed inner portion 120). Also, as shown in FIG. 4 , the inlet portion 126 of each cooling channel 124 may be defined in the additively printed inner partial cavity 144 to allow cooling fluid F to flow from the additively printed inner partial cavity 144 through the cooling channel 124 and be discharged from the airfoil 101 through the outlet portion 128 of each cooling channel 124.

[0052] 4 are defined at the same radial location as the respective cooling channels 124, it will be appreciated that the feed channels 138 need not be defined at the same radial location as any cooling channels 124 defined by the additively printed inner portion 120. For example, the number of feed channels 138 defined by the airfoil 101 may be different than the number of cooling channels 124 defined by the additively printed inner portion 120, such that not all feed channels 138 may be radially aligned with their respective cooling channels 124. For example, a fewer number of feed channels 138 may be defined in the airfoil 101 than the number of cooling channels 124 defined in the additively printed inner portion 120, and one or more of the feed channels 138 may be defined at a different radial location than the respective cooling channels 124. As another example, it may be desirable to vary the shape or geometry of the fluid path through the airfoil 101, e.g., to optimize heat transfer between the cooling fluid and the composite component, and so the radial location of the feed channel 138 may be different from the radial location of the respective cooling channel 124.

[0053] As described herein and as shown most clearly in FIG. 2A , in at least some embodiments, the outlets 128 of one or more cooling channels 124 of the additively printed inner portion 120 can be defined at an outer surface of the composite material component (e.g., at the cover outer surface 132, etc.). However, in some embodiments, the CMC outer portion 122 can overlie the outlets 128 of the one or more cooling channels 124. With reference to FIG. 4 , in such embodiments, exhaust channels 146 can be defined from the outer surface of the composite material component (e.g., from the cover outer surface 132) to the outlets 128 of each cooling channel 124, allowing the cooling fluid F to be exhausted from each cooling channel 124 to an environment external to the composite material component. 4 depicts exhaust channels 146 extending from the cooling channels 124 to the cover outer surface 132, it will be appreciated that in other embodiments, one or more cooling channels 124 of an additive printed inner portion 120 having an additive printed inner portion cavity 144 as shown in FIG. 4 can extend to the cover outer surface 132, with an outlet 128 of such cooling channel 124 defined in the cover outer surface 132, as described with respect to FIG. 2A. Additionally, although not depicted in FIG. 4, one or more exhaust channels 146 can also be defined from the cover outer surface 132 to the additive printed inner portion cavity 144. For example, cooling fluid F can be provided to the pressure side 102, the suction side 104, or both, via one or more exhaust channels 146 extending from the additively printed inner partial cavity 144 and / or the cooling channels 124 to the cover outer surface 132 on the pressure side 102, the suction side 104, or both.

[0054] In other embodiments, the additive printed inner portion 120 may be positioned elsewhere on the airfoil 101, i.e., the additive printed inner portion 120 need not be located adjacent the trailing edge 112 of the airfoil 101 (e.g., in the trailing edge region 136). For example, with reference to Figures 5 and 6, the additive printed inner portion 120 may be positioned adjacent the leading edge 110. As another example, with reference to Figure 7, the additive printed inner portion 120 may be positioned between the leading edge 110 and the trailing edge 112 along the axial direction A. For example, the additive printed inner portion 120 can be positioned along or adjacent to the suction side 104 between the cooling cavity 134 and the additive printed inner portion cavity 144, as shown in Figure 7, or can be positioned along or adjacent to the pressure side 102 between the cooling cavity 134 and the additive printed inner portion cavity 144. In the embodiments of Figures 5, 6, and 7, the inner portion 120 is enclosed within the CMC outer portion 122.

[0055] 5, 6, and 7, in some embodiments, the cooling channels 124 can be configured as an additively printed inner portion manifold 148. For example, rather than a plurality of cooling channels 124 forming a passage for cooling fluid F to flow through the additively printed inner portion 120 (e.g., from the inlet portion 126 to the outlet portion 128 of each cooling channel 124), the additively printed inner portion 120 can define an additively printed inner portion manifold 148. One or more feed channels 138 can deliver cooling fluid F from the cooling cavity 134 to the additively printed inner portion manifold 148, and one or more exhaust channels 146 can convey cooling fluid F from the additively printed inner portion manifold 148 to the cover outer surface 132 or to the exterior environment of the composite component (airfoil 101). Each feed channel 138 may extend from a feed channel inlet 140 defined in the cooling cavity 134 to a feed channel outlet 142 in an additively printed inner partial manifold 148. Each exhaust channel 146 may extend from an exhaust channel inlet 150 defined in the additively printed inner partial manifold 148 to an exhaust channel outlet 152 defined in the cover outer surface 132.

[0056] The additively printed inner partial manifold 148 can extend along the radial direction R (FIG. 2A) over all or a portion of the span S (FIG. 2A). For example, the additively printed inner partial manifold 148 can have a depth perpendicular to a length l defined between an inlet portion 126 of a cooling channel 124 defined at a radial location and an outlet portion 128 of the cooling channel 124 defined at the same radial location, for example as shown in FIG. 5, which is a cross-section of the airfoil 101 taken along a plane at a location along the span S (FIG. 2A). In some embodiments, a plurality of feed channels 138 can be defined at two or more radial locations and / or a plurality of exhaust channels 146 can be defined at two or more radial locations to deliver and / or convey cooling fluid F to and / or from the additively printed inner partial manifold 148 at different locations along the span S (FIG. 2A).

[0057] 6, the additively printed inner partial manifold 148 can be divided into a plurality of chambers 154 with ribs 156 disposed between and separating adjacent chambers 154. In some embodiments, the additively printed inner partial manifold 148 can be divided into two or more chambers 154 each extending radially, i.e., like the additively printed inner partial manifold 148, each chamber 154 extends across all or a portion of the span S (FIG. 2A) such that each chamber 154 has a depth. In other embodiments, the additively printed inner partial manifold 148 can be divided into two or more chambers 154 each extending generally along an axial direction A, i.e., a first chamber 154 can extend along a radial direction R over a first portion of the span S and a second chamber 154 can extend along a radial direction R over a second portion of the span S, such that the first chamber 154 and the second chamber 154 are defined at different radial locations rather than at different axial and / or circumferential locations.

[0058] 7, the exhaust channel 146 can be separated from the feed channel 138, for example, to optimize the thermal design. For example, a non-linear path can be defined from the feed channel inlet 140 to the exhaust channel outlet 152. By way of example, in the embodiment of FIG. 7, the feed channel 138 is defined in the airfoil 101 to feed the cooling fluid F to the additively printed inner portion 120, and the exhaust channel 146 is defined in the airfoil 101 to direct the cooling fluid F to the cover outer surface 132 of the airfoil 101 (or to the environment outside of the airfoil 101). The feed channel 138 and the exhaust channel 146 of the airfoil 101 shown in FIG. 7 are defined at the same radial location, but the feed channel 138 and the exhaust channel 146 are defined at an angle relative to the additively printed inner portion 120 and to the axial direction A, respectively. As such, the cooling fluid F must flow along a non-linear path from the feed channel inlet 140 to the exhaust channel outlet 152. For example, in the depicted embodiment, the cooling fluid F must flow generally aft to enter the additively printed inner portion 120 through the feed channel 138, then generally forward to exit the additively printed inner portion 120 through the exhaust channel 146, and finally generally aft to exit the exhaust channel 146 to the exterior of the airfoil 101.

[0059] The additive printed inner portion 120, feed channels 138, and exhaust channels 146 shown in the embodiments of Figures 5 and 6 similarly depict a non-linear path from the cooling cavity 134 through the additive printed inner portion 120 to the exterior of the airfoil 101, whereas the embodiments of Figures 2A, 2B, 3, and 4 depict a generally linear path for the cooling fluid F to flow from the cooling cavity 134 through the additive printed inner portion 120 to the exterior of the airfoil 101. The non-linear paths depicted in Figures 5 through 7, for example, may increase heat transfer or otherwise provide more beneficial cooling as compared to the linear paths through the additive printed inner portion 120 depicted. However, linear paths (such as those depicted in Figures 2A through 4, for example) may be desirable in some areas of the composite component to optimize cooling of such areas of the composite component.

[0060] Reference is now made to Figures 8A and 8B. Figure 8A provides a cross-sectional view of a composite material component 100 in a plane perpendicular to the span direction (or radial direction R, see Figure 8B), and Figure 8B provides a cross-sectional view of the composite material component 100 in a plane extending in the radial direction R and between the leading edge 110 and the trailing edge 112 (note that Figure 8B is a schematic plan view to depict certain aspects as discussed below). The composite material component 100 of Figures 8A and 8B can be configured in a manner similar to one or more of the exemplary composite material components described above.

[0061] 8A and 8B, one or more exhaust channels 146 can be defined in the additively printed inner portion cavity 144 for flow of cooling fluid F from the additively printed inner portion 120 to the cover outer surface 132 of the airfoil 101. For example, as shown in FIG. 8A, multiple exhaust channels 146 can be defined at multiple axial locations, for example, from one or more portions of a respective serpentine path 164 (see schematic plan view in FIG. 8B). In some embodiments, multiple exhaust channels 146 can be defined to a particular portion of the serpentine path 164, such as, for example, a first exhaust channel 146A, a second exhaust channel 146B, and a third exhaust channel 146C, which are defined to a first portion 170 of the first serpentine path 164A, for example, at or adjacent to the leading edge 110 (e.g., in the leading edge region 111). 8B , multiple exhaust channels 146 may be defined at approximately the same axial location but at different radial locations (e.g., at or adjacent to trailing edge 112). For example, as described with respect to FIG. 2B , in some embodiments, additive printed inner portion 120 may extend to trailing edge 112 and define multiple outlets 128, it being recognized that additive printed inner portion 120, rather than exhaust channels 146, directs cooling fluid F to cover outer surface 132 and / or to an environment external to airfoil 101.

[0062] In one or more of the embodiments discussed above (e.g., one or more of the embodiments of Figures 2A-8B), the additively printed inner portion 120 is formed from a first material and the CMC outer portion 122 is formed from a second material. However, in certain embodiments, one or more properties of the first and second materials are substantially similar or the same. For example, the first material can be a monolithic ceramic material (i.e., does not include fibers) and the second material can be a fiber-reinforced ceramic matrix composite (CMC), where the monolithic ceramic material is the same or substantially similar to the ceramic matrix of the fiber-reinforced CMC material. That is, in various embodiments, the CMC outer portion 122 can be formed from a second material that includes fibers disposed in a ceramic matrix, and the additively printed inner portion 120 can be formed from a first material that includes a ceramic matrix. In certain embodiments, the ceramic matrix of the additive printed inner portion 120 may be different from the ceramic matrix of the CMC outer portion 122, i.e., the first material may be different from the second material in ways other than fiber reinforcement, while in other embodiments, the ceramic matrix of the additive printed inner portion 120 may be the same as the ceramic matrix of the CMC outer portion 122. In either case, one or more properties of the first material (from which the additive printed inner portion 120 is formed) and the second material (from which the CMC outer portion 122 is formed) may be the same, such that the additive printed inner portion 120 and the CMC outer portion 122 are formed from materials that are compatible with respect to bonding, thermal expansion, etc. That is, the first and second materials may be selected with one or more common properties to provide sufficient bonding between the additive printed inner portion 120 and the CMC outer portion 122 to avoid component failure due to mismatch in thermal expansion coefficients, etc.

[0063] The additive printed inner portion 120 can be formed by a variety of methods. For example, the additive printed inner portion 120 can be a ceramic component (e.g., a silicon carbide (SiC) component, etc.) formed by an additive manufacturing process. In some embodiments, the additive printed inner portion 120 can be an additively formed SiC composite or an additively formed SiC monolith. In an exemplary embodiment, the additive printed inner portion 120 is formed using an additive manufacturing process. For example, the additive printed inner portion 120 includes multiple layers formed by depositing layers of additive material on a bed of an additive manufacturing machine and by selectively directing energy from an energy source onto the layers of additive material to melt a portion of the additive material. By such an additive process, the additive printed inner portion can be an integrally formed component (e.g., a single monolithic component, etc.). Additive manufacturing is described in more detail elsewhere herein.

[0064] 9 and 10, illustrations of composite components according to two additional embodiments are provided illustrating that the composite component need not be an airfoil (such as that shown in FIG. 2A), but can be another composite component (such as another composite component of turbofan engine 10 (FIG. 1)).

[0065] 9, a schematic diagram of a composite component is depicted. For example, in some embodiments, the composite component may be a nozzle airfoil assembly 100' having at least one of an inner band 174 or an outer band 176. The nozzle airfoil assembly 100' may also include an airfoil 101 extending from at least one of the inner band 174 or the outer band 176, it being recognized that the airfoil 101 may be configured as described elsewhere herein.

[0066] Continuing with FIG. 9, the nozzle airfoil assembly 100′ further includes an additive printed inner portion 120′ disposed in at least one of the inner band 174 or the outer band 176. In the embodiment of FIG. 9, the additive printed inner portion 120′ is disposed in the inner band 174, while in other embodiments the additive printed inner portion 120′ may be disposed in the outer band 176, or a first additive printed inner portion 120′ may be disposed in the inner band 174 and a second additive printed inner portion 120′ may be disposed in the outer band 176. The additive printed inner portion 120′ defines a cooling channel 124′ having an inlet portion 126′ and an outlet portion 128′. Additionally, the nozzle airfoil assembly 100' defines the cooling cavity 134', and the CMC outer portion 122' defining the cover outer surface 132' surrounds (encloses) the additively printed inner portion 120'. In FIG. 9 , the additively printed inner portion 120' is disposed in an aft portion of the inner band 174, for example, generally adjacent the trailing edge 112 of the airfoil 101 extending between the illustrated inner and outer bands 174, 176 of the nozzle airfoil assembly 100'. In other embodiments, the additively printed inner portion 120' can be disposed in another region of the inner band 174 (e.g., in a forward portion adjacent the leading edge 110 of the airfoil 101) or in a forward or aft portion of the outer band 176.

[0067] In the depicted embodiment, the inlet 126' of the cooling channel 124' is defined in the cooling cavity 134' and the exhaust channel 146' is defined from the exhaust channel inlet 150' and the outlet 128' of the cooling channel 124' to the exhaust channel outlet 152' at the cover outer surface 132'. In other embodiments, the inlet 126' of the cooling channel 124' can be defined in the cooling cavity 134' and the outlet 128' can be defined in the cover outer surface 132'. In yet other embodiments, a feed channel (not shown) (e.g., such as those described above) can be defined from a feed channel inlet in the cooling cavity 134' to a feed channel outlet at the inlet 126' of the cooling channel 124', and an exhaust channel 146' can be defined from the outlet 128' of the cooling channel 124' to the cover outer surface 132', or the outlet 128' can be defined in the cover outer surface 132'.

[0068] It will be appreciated that other variations and configurations of the additive printed inner portion 120 as described herein can be applied to the additive printed inner portion 120', and the additive printed inner portion 120' can be configured as described with respect to the additive printed inner portion 120. For example, the cooling channels 124' of the additive printed inner portion 120' can be configured as additive printed inner portion manifolds (such as those described with respect to the additive printed inner portion manifold 148 of FIGS. 5-7). As another example, the additive printed inner portion 120' can be configured with one or more serpentine paths (such as those described with respect to the serpentine paths 164 of FIGS. 8A and 8B).

[0069] 10, a schematic diagram of another composite component 100" is provided. In the depicted embodiment, the composite component 100 may be a shroud 100". For example, the shroud 100" may be positioned radially adjacent to a plurality of blades 178 (such as, for example, the HP turbine rotor blades 70 or the LP turbine rotor blades 74 of the turbofan engine 10 of FIG. 1). Similar to the airfoil 101 and the nozzle airfoil assembly 100', the shroud 100" includes an additively printed inner portion 120" having an outer surface 130" and defining at least one cooling channel 124" having an inlet portion 126" and an outlet portion 128". The shroud 100'' further includes a CMC outer portion 122'' having a cover outer surface 132'' and surrounding the additively printed inner portion 120'' such that the CMC outer portion 122'' is in contact with the outer surface 130''.

[0070] As shown in FIG. 10 , the additively printed inner portion 120″ is disposed in a portion of the shroud 100″ closest to the plurality of blades 178. In other embodiments, the additively printed inner portion 120″ may be positioned in other areas of the shroud 100″ (such as, for example, in a radially extending portion 180 that may be used to attach the shroud 100″ to a shroud hanger assembly). Additionally, while depicted in FIG. 10 as an open shroud 100″, it will be recognized that the subject matter described herein may also be applied to box shrouds or other shroud configurations.

[0071] As described elsewhere herein, the shroud 100" can define a cooling cavity 134" that receives a cooling fluid F. One or more feed channels 138" can be defined from the cooling cavity 134" to the additively printed inner portion 120" to feed the cooling fluid F to the cooling channels 124". Additionally or alternatively, an additively printed inner portion 120" can be positioned within the shroud 100" and the CMC outer portion 122" to provide a cooling An inlet 126" of the cooling channel 124" (or one or more inlets 126" of the cooling channel 124" for an additively printed inner portion 120" defining multiple cooling channels 124") is defined in the cooling cavity 134". Further, one or more exhaust channels 146" may be defined from the cooling channel 124" to the cover outer surface 132" to direct cooling fluid F from the additively printed inner portion 120" to an environment external to the shroud 100", for example, to provide cooling adjacent an aft portion of the blade 178. Additionally or alternatively, the additive printed inner portion 120" can be positioned within the shroud 100" and the CMC outer portion 122" such that an outlet portion 128" of the cooling channel 124" (or an outlet portion 128" of one or more cooling channels 124" for an additive printed inner portion 120" defining multiple cooling channels 124") is defined in the cover outer surface 132".

[0072] As with the nozzle airfoil assembly 100', it will be appreciated that other variations and configurations of the additive printed inner portion 120 as described herein can be applied to the additive printed inner portion 120", such that the additive printed inner portion 120" can be configured as described with respect to the additive printed inner portion 120. For example, the cooling channels 124" of the additive printed inner portion 120" can be configured as an additive printed inner portion manifold (such as those described with respect to the additive printed inner portion manifold 148 of FIGS. 5-7). As another example, the additive printed inner portion 120" can be configured with one or more serpentine paths (such as those described with respect to the serpentine path 164 of FIGS. 8A and 8B).

[0073] As shown in the various figures, in at least some embodiments, the additively printed inner portion 120 includes a plurality of heat transfer elements 182 that extend into at least one cooling channel 124. For example, with reference to Fig. 3, the additively printed inner portion 120 includes a plurality of heat transfer elements 182 that extend into the depicted cooling channel 124, and with reference to Fig. 7, the additively printed inner portion 120 includes a plurality of heat transfer elements 182 that extend into the depicted additively printed inner portion manifold 148. Similarly, the additive printed inner portion 120' shown in FIG. 9 includes a plurality of heat transfer elements 182' that extend into the depicted cooling channels 124', and in FIG. 10 the additive printed inner portion 120" includes a plurality of heat transfer elements 182" that extend into the cooling channels 124". The heat transfer elements 182 can, for example, increase heat transfer by the cooling fluid F flowing through the respective cooling channels 124 or the additive printed inner portion manifold 148. The heat transfer elements 182 can be configured as pins, fins, turbulators, etc.

[0074] 11A and 11B, in some embodiments, the additive printed inner portion 120, 120′, 120″ can be configured with a cellular structure 184 (e.g., a honeycomb cellular structure, etc.). The cellular structure 184 can provide for heat exchange between, for example, a cooling fluid F flowing through the additive printed inner portion 120 and the composite component (the airfoil 101 in the embodiment of FIGS. 11A and 11B) while also structurally reinforcing the composite component. For example, the orientation of the cellular structure 184 within the composite component can be used to directionally reinforce the strength of the composite component based on, for example, component loading. The orientation of the cellular structure 184 can be different for different composite components, for example, nozzle airfoils and blades can have directional strength differences, or the ... With respect to, for example, a nozzle airfoil immediately downstream of the combustion section 26 (FIG. 1) can have a directional strength difference compared to a nozzle airfoil disposed in the LP turbine 30 (FIG. 1). The cells 186 of the cell structure 184 can be oriented radially, axially, vertically, horizontally, or in any orientation in a two-dimensional or three-dimensional pattern. Additionally, in some embodiments, the cells 186 can be closed and isolated from one another, and in alternative embodiments, the cells 186 can be open and connected to one another. In still other embodiments, the cell structure 184 can be a hybrid of open and connected cells 186 and closed and isolated cells 186, e.g., in one portion of the cell structure 184, the cells 186 are open and connected to one another, and in another portion of the cell structure 184, the cells 186 are closed and isolated from one another.

[0075] 12, the additively printed inner portion 120 can be utilized as a mandrel during the manufacturing process of the resulting composite material component 100, which is then retained in the part as it is formed. The additively printed inner portion 120 allows complex mandrel shapes to be utilized. As such, a non-line of sign mandrel can be utilized without subsequent extraction from the part after consolidation of the resulting composite material component 100. Additionally, multiple cooling cavities 134 can be defined in the additively printed inner portion 120.

[0076] The outer surface 130 of the additive-printed inner portion 120 generally serves as a mandrel for the formation of the CMC outer portion 122. For example, one or more composite plies can be wrapped substantially or completely around the additive-printed inner portion 120 to form the CMC outer portion 122. Although shown to include a single additive-printed inner portion 120 as the mandrel, it should be understood that any number of additive-printed inner portions 120 can be included.

[0077] 13, a flow diagram of a method 200 for forming a composite material component according to an exemplary embodiment of the present disclosure is provided. The method 200 of FIG. 13 may be utilized to form one or more of the exemplary composite material components (e.g., airfoil 100, nozzle airfoil assembly 100′, and shroud 100″, etc.) described above with reference to FIGS. 2A-11B. Thus, it will be appreciated that the method 200 may be utilized to form a composite material component that generally includes an additively printed inner portion surrounded by a CMC outer portion. However, in other exemplary embodiments, the method 200 may additionally or alternatively be utilized to form any other suitable composite material component.

[0078] As depicted, method 200 includes forming an insert (e.g., an additively printed inner portion) at 202 that defines at least one cooling channel having an inlet and an outlet. As described herein, the insert can be formed by an additive manufacturing process, although in some embodiments, other processes can be used as well. For an additively printed inner portion formed by additive manufacturing, method 200 further includes depositing a layer of additive material on a bed of an additive manufacturing machine at 204, and selectively directing energy from an energy source onto the layer of additive material at 206 to melt a portion of the additive material and form the additively printed inner portion. By utilizing such an additive manufacturing process, the additively printed inner portion can be formed with one or more cooling channels as described herein, where the one or more cooling channels have any desired location, shape, size, characteristics, etc. For example, a plurality of heat transfer elements (e.g., pins, fins, or turbulators, etc.) can be formed into one or more cooling channels when a layer of additive material is deposited onto the bed of an additive manufacturing machine, such that when the additive material is melted as described herein, the additively printed inner portion includes at least one cooling channel having a plurality of heat transfer elements formed therein.

[0079] Continuing with FIG. 13, the method 200 also includes disposing an insert in the CMC outer portion having a cover outer surface, at 208. For example, one or more composite plies can be wrapped around the additively printed inner portion to form the CMC outer portion, or the CMC outer portion can be a composite preform and the additively printed inner portion is additively printed or positioned into the composite preform. The CMC outer portion can then be formed by chemical vapor infiltration (CVI), melt infiltration (MI), polymer impregnation and pyrolysis (PIP) of the preform or multiple plies (e.g., with silicon) surrounding the inner portion.

[0080] As described herein, the CMC outer portion substantially surrounds the insert such that the CMC outer portion is in contact with the outer surface. Also, as described herein, the composite material component (e.g., the CMC outer portion) can define a cooling cavity for receiving a cooling fluid. In some embodiments, the insert can be disposed in the CMC outer portion such that an inlet of at least one cooling channel is disposed in the cooling cavity for the cooling fluid to flow from the cooling cavity into the at least one cooling channel. In other embodiments, the CMC outer portion of the composite material component or other material is disposed between the insert and the cooling cavity, and the method 200 optionally includes, at 210, defining one or more feed channels from the cooling cavity to the at least one cooling channel to provide a passageway for the cooling fluid to flow from the cooling cavity into the at least one cooling channel. The one or more feed channels can be defined by laser drilling, electrical discharge machining (EDM), or any other suitable process for forming a passageway from the cooling cavity to the additively printed inner portion.

[0081] Further, in some embodiments, the insert can be disposed within the CMC outer portion such that an outlet of the at least one cooling channel extends to the cover outer surface. In other embodiments, the CMC outer portion of the composite component or other material is positioned over the outlet of the at least one cooling channel or between the inserts at the cover outer surface, and the method 200 optionally includes, at 212, defining one or more exhaust channels from the cover outer surface to the at least one cooling channel of the insert to provide a passageway for cooling fluid flowing into the at least one cooling channel to exit the cooling channel and exit the composite component at the cover outer surface. The one or more exhaust channels can be defined by laser drilling, EDM, or any other suitable process for forming a passageway from the cover outer surface to the insert.

[0082] While the various composite material components 100, 100', 100" described and illustrated herein each incorporate an insert (e.g., a single additively printed inner portion 120, 120', 120"), it will be appreciated that in some embodiments, the composite material component 100, 100', 100" may include multiple inserts (e.g., additively printed inner portions 120, 120', 120"). For example, if the composite material component is an airfoil 101, the airfoil 101 may include a first additively printed inner portion 120 in the trailing edge region 136 of the airfoil 101 and a second additively printed inner portion 120 positioned near the leading edge 110 of the airfoil 101. Thus, the composite material component 100, 100', 100" may include one, two, three, or more separate additively printed inner portions 120, 120', 120". Each additively printed inner portion 120, 120', 120" may be positioned within each composite component 100, 100', 100" at or near where cooling is desired.

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

[0084] 1. A composite component for a gas turbine engine, the composite component comprising: an additive-printed inner portion defining at least one flow path feature, the additive-printed inner portion comprising SiC; and a ceramic matrix composite (CMC) outer portion formed on the additive-printed inner portion such that the CMC outer portion substantially surrounds the additive-printed inner portion, the CMC outer portion comprising a fiber reinforced ceramic matrix and defining at least one cooling cavity fluidly coupled to the at least one flow path feature of the additive-printed inner portion, the fiber reinforced ceramic matrix comprising SiC.

[0085] 13. The composite component of any preceding claim, wherein at least one flow path feature of the additively printed inner portion has a non-linear geometry.

[0086] 2. The composite component of any preceding claim, wherein the CMC outer portion defines a linear channel fluidly connecting the at least one flow path feature and the at least one cooling cavity.

[0087] 2. The composite component of any preceding claim, wherein the CMC outer portion defines a hollow airfoil having a cooling cavity therein, and wherein at least one flow path feature extends nonlinearly from the cooling cavity to an exterior surface of the hollow airfoil.

[0088] 13. The composite component of any preceding claim, wherein the hollow airfoil has a pressure side opposite the suction side and a leading edge opposite the trailing edge, the pressure side and suction side extending axially between the leading and trailing edges with a nonlinear curved geometry and extending radially over a predetermined span, the cooling cavity is positioned adjacent the leading edge, and the at least one flow path feature includes a plurality of cooling passages spaced apart along the predetermined span, the plurality of cooling passages each extending adjacent the suction side from an inlet in the cooling cavity to an outlet at the trailing edge with a nonlinear curved geometry in the axial direction that matches the nonlinear curved geometry of the suction side.

[0089] 2. The composite component of any preceding claim, wherein at least one flow path feature is a cooling channel having a primary flow direction and an interior surface defined by a plurality of heat transfer elements extending into the cooling channel non-parallel to the primary flow direction.

[0090] 13. The composite component of any preceding claim, wherein the CMC outer portion defines an airfoil and the additively printed inner portion is positioned in a trailing edge region of the airfoil.

[0091] 13. The composite component of any preceding claim, wherein the CMC outer portion defines an airfoil and the additively printed inner portion is positioned in a leading edge region of the airfoil.

[0092] 2. The composite component of any preceding claim, wherein the CMC outer portion defines an airfoil and the additively printed inner portion is positioned adjacent a suction side of the airfoil.

[0093] 2. The composite component of any preceding claim, wherein the CMC outer portion defines a nozzle airfoil assembly including an inner band and an outer band, and the additively printed inner portion is positioned within at least one of the inner band or the outer band.

[0094] 2. The composite component of any preceding claim, wherein the CMC outer portion defines a shroud including an outer surface positioned radially adjacent the plurality of rotatable blades, and the additively printed inner portion is positioned adjacent the outer surface.

[0095] 13. The composite component of any preceding paragraph, wherein the CMC outer portion is formed by chemical vapor infiltration of a preform or multiple plies that are laminated over an additively printed inner portion.

[0096] 13. The composite component of any preceding claim, wherein the CMC outer portion is formed by melt infiltration of a preform or multiple plies that are layered over an additively printed inner portion.

[0097] 13. The composite component of any preceding claim, wherein the CMC outer portion surrounds the additively printed inner portion.

[0098] 1. A composite component for a gas turbine engine, the composite component including: an additive-printed inner portion defining at least one cooling channel having an inlet and an outlet, the additive-printed inner portion having an outer surface; and a ceramic matrix composite (CMC) outer portion substantially surrounding the additive-printed inner portion such that the CMC outer portion is in contact with the outer surface, the CMC outer portion having a cover outer surface, an outlet portion of the at least one cooling channel extending to the cover outer surface.

[0099] 2. The composite component of any preceding claim, wherein the CMC outer portion defines a hollow airfoil having a cooling cavity therein, and the at least one flow path feature extends from the cooling cavity to a cover outer surface of the CMC outer portion.

[0100] 2. The composite component of any preceding claim, wherein the additively printed inner portion comprises SiC and the CMC outer portion comprises a fiber reinforced ceramic matrix, the ceramic matrix comprising SiC.

[0101] 2. The composite component of any preceding claim, wherein the CMC outer portion defines a hollow airfoil, and the at least one flow path feature includes a cooling cavity, and the at least one flow path feature extends nonlinearly from the cooling cavity to an exterior surface of the hollow airfoil.

[0102] 13. The composite component of any preceding claim, wherein the hollow airfoil has a pressure side opposite the suction side and a leading edge opposite the trailing edge, the pressure side and the suction side extending axially between the leading and trailing edges with a nonlinear curved geometry and extending radially for a predetermined span, the at least one cooling channel is positioned adjacent the leading edge, and the at least one flow path feature includes a plurality of cooling passages spaced apart along the predetermined span, the plurality of cooling passages each extending adjacent the suction side from an inlet in the at least one cooling channel to an outlet at the trailing edge with a nonlinear curved geometry along the axial direction that matches the nonlinear curved geometry of the suction side.

[0103] 13. The composite material component of any preceding claim, wherein the additively printed inner portion includes a plurality of layers formed by depositing layers of additive material on a bed of an additive manufacturing machine and by selectively directing energy from an energy source onto the layers of additive material to melt portions of the additive material, wherein the additively printed inner portion is integrally formed as a single monolithic component.

[0104] 1. A composite material component comprising: an insert defining at least one cooling channel having an inlet portion and an outlet portion, the insert having an insert outer surface; and a composite material cover surrounding the insert such that the composite material cover is in contact with the insert outer surface, the composite material cover having a cover outer surface, wherein an outlet portion of the at least one cooling channel extends to the cover outer surface.

[0105] 2. The composite component of any preceding claim, wherein the insert is formed from a first material and the composite cover is formed from a second material, and wherein one or more properties of the first material and the second material are the same.

[0106] 2. The composite component of any preceding claim, wherein the first material is a monolithic ceramic material and the second material is a fiber reinforced ceramic matrix composite.

[0107] 13. The composite component of any preceding claim, wherein the composite component is an airfoil having a pressure side opposite the suction side and a leading edge opposite the trailing edge, the pressure side and the suction side extending axially between the leading edge and the trailing edge and extending radially for a span, and the at least one cooling channel is a plurality of cooling channels, each cooling channel of the plurality of cooling channels having an inlet portion and an outlet portion, the outlet portion of each cooling channel being defined in the cover outer surface adjacent the trailing edge.

[0108] 13. The composite component of any preceding claim, wherein the composite component further includes a component cavity for receiving a cooling fluid, the trailing edge being defined by a trailing edge portion positioned aft of the component cavity, and the insert being disposed in the trailing edge portion such that an axially extending midline of the insert is closer to one of the pressure side or suction side than to the other of the pressure side or suction side.

[0109] 2. The composite component of any preceding claim, further comprising a component cavity for receiving a cooling fluid, the insert being positioned at a rear of the component cavity, and the feed channel extending from the component cavity to an inlet of each of the plurality of cooling channels.

[0110] The composite material component of any preceding claim, further comprising a component cavity for receiving a cooling fluid, the insert being positioned at a rear of the component cavity, the insert defining an insert cavity extending along at least a portion of the predetermined span, and at least one feed channel extending from the component cavity to the insert cavity.

[0111] 13. The composite component of any preceding claim, further comprising a component cavity for receipt of a cooling fluid, the feed channel extending from the component cavity to an inlet of at least one cooling channel.

[0112] 13. The composite component of any preceding claim, further comprising a component cavity for receipt of a cooling fluid, the composite component being an airfoil having a pressure side opposite the suction side and a leading edge opposite the trailing edge, the pressure and suction sides extending axially between the leading and trailing edges and extending radially for a span, and the insert being positioned adjacent the leading edge.

[0113] 13. The composite component of any preceding claim, further comprising a component cavity for receipt of a cooling fluid, the composite component being an airfoil having a pressure side opposite the suction side and a leading edge opposite the trailing edge, the pressure and suction sides extending axially between the leading and trailing edges and extending radially for a span, the insert being positioned axially between the leading and trailing edges, the insert being positioned adjacent one of the pressure and suction sides.

[0114] 2. The composite material component of any preceding claim, further comprising a component cavity for receiving a cooling fluid, a feed channel, and an exhaust channel, wherein at least one cooling channel is configured as an insert manifold for receiving a cooling fluid, the insert manifold having a depth that is perpendicular to a length defined between an inlet portion of the at least one cooling channel and an outlet portion of the at least one cooling channel, the feed channel extending from a feed channel inlet portion defined in the component cavity to a feed channel outlet portion in the insert manifold, and the exhaust channel extending from a exhaust channel inlet portion defined in the insert manifold to an exhaust channel outlet portion at the cover outer surface.

[0115] 2. The composite material component of any preceding claim, wherein the composite material component extends along a radial direction between an inner end and an outer end, the insert extends along a radial direction from the inner end to the outer end, the insert extends along a direction perpendicular to the radial direction between a first end and a second end, the insert defines a serpentine cavity for receiving a cooling fluid, the serpentine cavity having a cavity inlet adjacent one of the first end or the second end, and the serpentine cavity includes at least a first directional change adjacent one of the inner end or the outer end and a second directional change adjacent the other of the inner end or the outer end.

[0116] 13. The composite component of any preceding claim, wherein the composite component is a nozzle airfoil assembly including at least one of an inner band or an outer band, and the insert is disposed within at least one of the inner band or the outer band.

[0117] 13. The composite component of any preceding claim, wherein the nozzle airfoil assembly defines a component cavity, and an inlet of the at least one cooling channel is defined in the component cavity.

[0118] 13. The composite component of any preceding claim, wherein the composite component is a shroud positioned radially adjacent to a plurality of blades.

[0119] 13. The composite component of any preceding claim, wherein the insert defines a plurality of cooling channels, each of the plurality of cooling channels having an inlet portion and an outlet portion, the outlet portion of each of the plurality of cooling channels being defined in the cover outer surface.

[0120] 13. The composite component of any preceding claim, wherein the insert includes a plurality of heat transfer elements extending into the at least one cooling channel.

[0121] 2. The composite material component of any preceding claim, wherein at least one cooling channel is configured as an insert manifold for receiving a cooling fluid, the insert manifold having a depth that is perpendicular to a length defined between an inlet portion of the at least one cooling channel and an outlet portion of the at least one cooling channel, and the insert includes a plurality of heat transfer elements extending into the insert manifold.

[0122] 2. The composite material component of any preceding claim, wherein the insert includes a plurality of layers formed by depositing layers of additive material on a bed of an additive manufacturing machine and selectively directing energy from an energy source onto the layers of additive material to melt portions of the additive material, and wherein the insert is integrally formed as a single monolithic component.

[0123] 13. The composite component of any preceding claim, wherein the CMC outer portion surrounds the additively printed inner portion.

[0124] 1. A method for forming a composite material component for a gas turbine engine, the method including: substantially enclosing an additively printed inner portion in a preform or a plurality of plies that are laminated over the additively printed inner portion; and infiltrating the preform or the plurality of plies with silicon via chemical vapor infiltration or melt infiltration to form a CMC outer portion having a covering outer surface.

[0125] 13. The method of any preceding claim, wherein the additively printed inner portion defines at least one cooling channel having an inlet and an outlet.

[0126] 13. The method of any preceding claim, wherein an outlet of the at least one cooling channel is in fluid communication through an exterior surface of the cover.

[0127] 1. A method for forming a composite material component, the method including the steps of forming an insert defining at least one cooling channel having an inlet portion and an outlet portion, the insert having an insert outer surface, and disposing the insert within a composite material cover having a cover outer surface, the composite material cover surrounding the insert such that the composite material cover is in contact with the insert outer surface, the outlet portion of the at least one cooling channel extending to the cover outer surface.

[0128] 2. The method of any preceding claim, wherein the CMC outer portion surrounds the additively printed inner portion.

[0129] This written description uses examples to disclose the disclosure, including the best mode, and to enable any person skilled in the art to practice the disclosure, including making and using any device or system and performing any incorporated methods. The patentable scope of the disclosure is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.

[0130] Further aspects of the invention are provided by the subject matter of the following clauses.

[0131] [Item 1] A composite material component for a gas turbine engine, the composite material component comprising: an additive-printed inner portion defining at least one flow path feature, the additive-printed inner portion comprising SiC; a ceramic matrix composite (CMC) outer portion, the CMC outer portion being formed over the additively printed inner portion such that the CMC outer portion substantially surrounds the additively printed inner portion; Including, A composite material component, wherein the CMC outer portion includes a fiber reinforced ceramic matrix and defines at least one cooling cavity fluidly coupled to at least one flow path feature of the additively printed inner portion, the fiber reinforced ceramic matrix including SiC.

[0132] [Item 2] A composite material component as described in any preceding item, wherein at least one flow path feature of the additively printed inner portion has a nonlinear geometric shape.

[0133] [Item 3] A composite component as described in any preceding item, wherein the CMC outer portion defines a linear channel fluidly connecting at least one flow path feature and at least one cooling cavity.

[0134] [Item 4] The composite component of any preceding item, wherein the CMC outer portion defines a hollow airfoil having a cooling cavity therein, and at least one flow path feature extends nonlinearly from the cooling cavity to an exterior surface of the hollow airfoil.

[0135] [Item 5] The composite component of any preceding item, wherein the hollow airfoil has a pressure side opposite the suction side and a leading edge opposite the trailing edge, the pressure side and the suction side extending axially between the leading edge and the trailing edge with a nonlinear curved geometry and extending radially for a predetermined span, the cooling cavity is positioned adjacent the leading edge, and the at least one flow path feature includes a plurality of cooling passages spaced apart along the predetermined span, the plurality of cooling passages each extending adjacent the suction side from an inlet in the cooling cavity to an outlet at the trailing edge with a nonlinear curved geometry along the axial direction that matches the nonlinear curved geometry of the suction side.

[0136] [Item 6] The composite component of any preceding item, wherein at least one flow path feature is a cooling channel having a primary flow direction and an interior surface defined by a plurality of heat transfer elements extending into the cooling channel non-parallel to the primary flow direction.

[0137] [Item 7] The composite component of any preceding item, wherein the CMC outer portion defines an airfoil and the additively printed inner portion is positioned within a trailing edge region of the airfoil.

[0138] [Item 8] The composite component of any preceding item, wherein the CMC outer portion defines an airfoil and the additively printed inner portion is positioned within a leading edge region of the airfoil.

[0139] [Item 9] The composite component of any preceding item, wherein the CMC outer portion defines an airfoil and the additively printed inner portion is positioned adjacent to a suction side of the airfoil.

[0140] [Item 10] The composite component of any preceding item, wherein the CMC outer portion defines a nozzle airfoil assembly including an inner band and an outer band, and the additively printed inner portion is positioned within at least one of the inner band or the outer band.

[0141] [Item 11] The composite component of any preceding item, wherein the CMC outer portion defines a shroud including an outer surface positioned radially adjacent to the plurality of rotatable blades, and the additively printed inner portion is positioned adjacent to the outer surface.

[0142] [Item 12] The composite component of any preceding item, wherein the CMC outer portion is formed by chemical vapor infiltration of a preform or multiple plies that are laminated over an additively printed inner portion.

[0143] [Item 13] The composite component of any preceding item, wherein the CMC outer section is formed by melt infiltration of a preform or multiple plies that are layered over an additively printed inner section.

[0144] [Item 14] The composite component of any preceding item, wherein the CMC outer portion surrounds the additively printed inner portion.

[0145] [Item 15] A composite component for a gas turbine engine, the composite component comprising: an additive printed inner portion defining at least one cooling channel having an inlet and an outlet, the additive printed inner portion having an outer surface; a ceramic matrix composite (CMC) outer portion substantially surrounding the additively printed inner portion such that the CMC outer portion is in contact with the outer surface, the CMC outer portion having a cover outer surface; Including, The composite component, wherein an outlet of the at least one cooling channel extends to an outer surface of the cover.

[0146] [Item 16] The composite component of any preceding item, wherein the CMC outer portion defines a hollow airfoil having a cooling cavity therein, and at least one flow path feature extends from the cooling cavity to a cover outer surface of the CMC outer portion.

[0147] [Item 17] The composite component of any preceding item, wherein the additively printed inner portion comprises SiC and the CMC outer portion comprises a fiber reinforced ceramic matrix, the fiber reinforced ceramic matrix comprising SiC.

[0148] [Item 18] The composite component of any preceding item, wherein the CMC outer portion defines a hollow airfoil, the at least one flow path feature includes at least one cooling channel, and the at least one flow path feature extends nonlinearly from the at least one cooling channel to an exterior surface of the hollow airfoil.

[0149] [Item 19] The composite component of any preceding item, wherein the hollow airfoil has a pressure side opposite the suction side and a leading edge opposite the trailing edge, the pressure side and the suction side extending axially between the leading edge and the trailing edge with a nonlinear curved geometry and extending radially for a span, and the at least one cooling channel is positioned adjacent the leading edge, and the at least one flow path feature includes a plurality of cooling passages spaced apart along the span, the plurality of cooling passages each extending adjacent the suction side from an inlet in the at least one cooling channel to an outlet at the trailing edge with a nonlinear curved geometry along the axial direction that matches the nonlinear curved geometry of the suction side.

[0150] [Item 20] A method for forming a composite material component for a gas turbine engine, the method comprising: substantially enclosing the additively printed inner portion within a preform or a plurality of plies that are laminated onto the additively printed inner portion; infiltrating the preform or the plurality of plies with silicon via chemical vapor infiltration or melt infiltration to form a CMC outer portion having a cover outer surface; A method comprising: [Explanation of symbols]

[0151] 10. Turbofan engine 12 Longitudinal centerline 14. Fan Section 16 Core Turbine Engine 18 Outer casing 20 Entrance 22 Booster or low pressure (LP) compressor 24 High Pressure (HP) Compressor 26 Combustion Section 28 High Pressure (HP) Turbine 30 Low Pressure (LP) Turbine 32 Jet Exhaust Nozzle Section 34 High Pressure (HP) Shaft or Spool 36 Low Pressure (LP) Shaft or Spool 38 Fans 40 Fan Blade 42 Discs 48 Front nacelle 50 Fan Case 52 Exit guide vane 54 Downstream Section 56 Bypass air flow passage 58 Air 60 Entrance 62 First Part of Air 64 Second Part of Air 66 Combustion Gas 68 HP Turbine Stator Vane 70 HP Turbine Rotor Blades 72 LP turbine stator vanes 74 LP turbine rotor blades 76 Fan nozzle exhaust section 78 Hot Gas Path 80 Compressor stage 82 Compressor vanes, compressor stator vanes 84 Compressor rotating blades, compressor rotor blades 90 Compressor casing 100, 100', 100” Composite Components 101 Airfoil 102 Pressure side 104 Negative pressure side 106 Inner end 108 Outer edge 110 Leading edge 111 Leading edge region 112 Trailing edge 120, 120', 120" printed interior 122, 122', 122” CMC outer section 124, 124', 124” cooling channel, flow path features 126, 126', 126" Entrance 128, 128', 128” exit section 130, 130” outer surface 132, 132', 132” Cover outer surface 134, 134', 134" Cooling Cavity 136 Posterior edge region 138, 138” Feed Channel 140 Feed channel inlet 142 Feed channel outlet 144 Inner Part Cavity 146, 146', 146” Discharge Channel 146A First discharge channel 146B Secondary discharge channel 146C Third discharge channel 148 Inner Part Manifold 150 Exhaust channel inlet 152 Discharge channel outlet 154 Chamber 156 Rib 164 Tortuous Route 164A First Serpentine Path 170 first portion of first serpentine path 164A 174 Inner Band 176 Outer Band 178 Blade 180 Radial extending portion 182, 182' Heat transfer element 184 Cell Structure 186 cells A Axial direction C Circumferential direction F Cooling fluid l Length M median line R Radial direction S span

Claims

1. A composite material component (100) for a gas turbine engine (10), said composite material component (100) comprising: an additive-printed inner portion (120) defining at least one flow path feature (124), said additive-printed inner portion (120) comprising SiC; a ceramic matrix composite (CMC) outer portion (122), said CMC outer portion (122) being formed on said additive-printed inner portion (120) such that said CMC outer portion (122) substantially surrounds said additive-printed inner portion (120); Including, the CMC outer portion (122) includes a fiber reinforced ceramic matrix and defines at least one cooling cavity fluidly coupled to at least one of the flow path features (124) of the additively printed inner portion (120); The composite component (100), wherein the fiber reinforced ceramic matrix comprises SiC.

2. The composite material component (100) of claim 1 , wherein at least one of the flow path features (124) of the additively printed inner portion (120) has a non-linear geometry.

3. 3. The composite material component (100) of claim 1 or 2, wherein the CMC outer portion (122) defines a linear channel fluidly connecting at least one of the flow path features (124) and at least one of the cooling cavities.

4. 4. The composite material component (100) of claim 1, wherein the CMC outer portion (122) defines a hollow airfoil (101) having a cooling cavity therein, and wherein at least one of the flow path features (124) extends nonlinearly from the cooling cavity to an exterior surface of the hollow airfoil (101).

5. 5. The composite component of claim 4, wherein the hollow airfoil has a pressure side opposite a suction side and a leading edge opposite a trailing edge, the pressure side and the suction side extending axially between the leading and trailing edges with a non-linear curved geometry and extending radially for a span, the cooling cavity positioned adjacent the leading edge, and the at least one flow path feature includes a plurality of cooling passages spaced apart along the span, each of the plurality of cooling passages extending adjacent the suction side from an inlet in the cooling cavity to an outlet at the trailing edge with a non-linear curved geometry along the axial direction that matches the non-linear curved geometry of the suction side.

6. 6. The composite material component of claim 1, wherein at least one of the flow path features is a cooling channel having a primary flow direction and an interior surface defined by a plurality of heat transfer elements extending into the cooling channel non-parallel to the primary flow direction.

7. 7. The composite material component (100) of claim 1, wherein the CMC outer portion (122) defines an airfoil (101), and the inner portion (120) is positioned in a trailing edge region of the airfoil (101).

8. 8. The composite material component (100) of claim 1, wherein the CMC outer portion (122) defines an airfoil (101) and the inner portion (120) is positioned in a leading edge region of the airfoil (101).

9. 9. The composite material component (100) of any one of claims 1 to 8, wherein the CMC outer portion (122) defines an airfoil (101) and the inner portion (120) is positioned adjacent a suction side (104) of the airfoil (101).

10. 10. The composite material component of claim 1, wherein the CMC outer portion defines a nozzle airfoil assembly including an inner band and an outer band, the inner portion being positioned within at least one of the inner band or the outer band.

11. 11. The composite material component (100) of claim 1, wherein the CMC outer portion (122) defines a shroud including an outer surface positioned radially adjacent a plurality of rotatable blades, and the inner portion (120) is positioned adjacent the outer surface.

12. the CMC outer portion (122) is formed by chemical vapor infiltration of a preform or multiple plies that are layered onto the inner portion (120); or 12. The composite material component (100) of claim 1, wherein the CMC outer portion (122) is formed by melt infiltration of a preform or multiple plies that are layered over the inner portion (120).

13. The composite material component (100) of any one of claims 1 to 12, wherein the CMC outer portion (122) surrounds the additively printed inner portion (120).

14. A composite material component (100) for a gas turbine engine (10), the composite material component (100) comprising: an additive printed inner portion (120) defining at least one cooling channel having an inlet and an outlet, said additive printed inner portion (120) having an outer surface; a CMC outer portion (122), said CMC outer portion (122) substantially surrounding said additively printed inner portion (120), said CMC outer portion (122) being in contact with said outer surface, said CMC outer portion (122) having a cover outer surface; Including, The outlet of at least one of the cooling channels extends to an outer surface of the cover.

15. A method for forming a composite material component (100) for a gas turbine engine (10), the method comprising: Substantially enclosing the additively printed inner portion (120) in a preform or a number of plies that are laminated onto said inner portion (120); infiltrating the preform or the plurality of plies with silicon via chemical vapor infiltration or melt infiltration to form a CMC outer portion (122) having a cover outer surface; A method comprising:

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