Internal Tooling for Composite Parts

Abstract

To provide systems and methods for composite part fabrication.SOLUTION: The method of fabricating a composite part comprises the steps of: selecting a tool (100) with sides (120, 130, 140) made of a core material in a desired size and shape; and disposing a preform (350) of fiber reinforced material that surrounds the tool, resulting in a lamina assembly (300) comprising the preform and the tool; heating the tool and the preform; co-bonding the tool to the fiber reinforced material within the lamina assembly; and hardening the preform and the tool into a hybrid composite part.SELECTED DRAWING: Figure 3

Description

[Technical field]

[0001] FIELD OF THE DISCLOSURE This disclosure relates to the field of manufacturing, and more particularly to composite parts. [Background technology]

[0002] Composite parts, such as carbon fiber reinforced polymer (CFRP) parts, are utilized to provide structural strength to vehicles and structures. During manufacturing, composite stringers or other structural supports can be formed by laying up sheets of fiber into a desired shape. This can result in a "C" shaped cross section, an "I" shaped cross section, a hat shaped cross section, or other shaped cross sections. When the sheets of fiber are laid up together and contoured to form a desired shape, gaps may form at the bend radii and / or the joints between the sheets. A challenge remains in filling these gaps, as it is undesirable to have gaps in the finished structure.

[0003] It would therefore be desirable to have a method and apparatus that takes into account at least some of the above-mentioned problems, as well as other possible problems. Summary of the Invention

[0004] The embodiments described herein provide tooling that defines the shape of the preform during layup and that is integrated with the preform during its consolidation into a hybrid composite part. Thus, while the tooling is originally used to define the shape for laying up the unconsolidated fiber reinforcement, the tooling is co-bonded with the fiber reinforcement during the consolidation process. This results in a hybrid composite part made from both the tooling and the fiber reinforcement. The tooling forms the core of the hybrid composite part and structurally reinforces the composite material within the hybrid composite part. Thus, the tooling acts as both a forming tool for the hybrid composite part and a structural strength enhancing part. This provides benefits in the form of increasing the ease of manufacture of the composite part as well as increasing its strength.

[0005] One embodiment is a method for manufacturing a composite part that includes the steps of selecting a tool having a surface shaped to a desired dimension from a core material, placing a preform of fiber reinforcement surrounding the tool to result in a laminar assembly including the preform and the tool, heating the tool and preform to co-bond the tool to the fiber reinforcement within the laminar assembly, and solidifying the preform and tool to form a hybrid composite part.

[0006] A further embodiment is a non-transitory computer readable medium embodied with program instructions operable, when executed by a processor, to perform a method of manufacturing a composite part, the method including selecting a tool having a surface shaped to a desired dimension from a core material, placing a preform of fiber reinforcement surrounding the tool to result in a laminar assembly including the preform and the tool, heating the tool and preform to co-bond the tool to the fiber reinforcement within the laminar assembly, and solidifying the preform and tool to form a hybrid composite part.

[0007] An additional embodiment is an article in the form of a hybrid composite part, the article including a tool and a fiber reinforcement surrounding and co-bonded with the tool, the tool surrounding a void.

[0008] Yet another embodiment is an apparatus in the form of a hybrid composite part. The apparatus includes tooling having a body with faces fabricated to a desired dimension from a core material and is made of a material that maintains rigidity at the cure temperature of the fiber reinforcement. The tooling also includes one or more flanges disposed at the location where the faces are bonded together. The hybrid composite part also includes a void within the body that runs along the axial length of the body.

[0009] The device, the medium, and the method are also referred to in the appendix below, which should not be confused with the claims.

[0010] 1. A method (200) for manufacturing a composite part, comprising: Selecting (202) a tool having a surface made to the desired dimensions from a core material; placing a preform of fiber reinforcement surrounding the tool, resulting in a lamination assembly comprising the preform and the tool (204); heating (206) the tool and the preform; co-bonding (208) the tool to the fiber reinforcement within the lamination assembly; consolidating the preform and the tool to form a hybrid composite part (210); The method includes:

[0011] 2. The method of claim 1, further comprising forming the tool such that a body of the tool includes a void.

[0012] 3. Fabricating the tool from a sheet of metal; 3. The method of claim 1 or 2, further comprising:

[0013] 4. Fabricating the tool via additive manufacturing; 4. The method of any one of claims 1 to 3, further comprising:

[0014] 5. transferring stresses during service of the hybrid composite component from a fiber reinforced portion of the hybrid composite component to the tool; 5. The method of any one of claims 1 to 4, further comprising:

[0015] 6. elastically deforming the hybrid composite component in response to an applied stress while maintaining a bond between the tool and a fiber reinforced portion of the hybrid composite component; 6. The method of any one of claims 1 to 5, further comprising:

[0016] 7. An aircraft part assembled in accordance with the method set out in any of appendixes 1 to 6.

[0017] According to a further aspect of the present medium, the following is provided:

[0018] 8. A non-transitory computer readable medium embodying program instructions operable, when executed by a processor, to perform a method of manufacturing a composite part, the method comprising: Selecting (202) a tool having a surface made to the desired dimensions from a core material; placing a preform of fiber reinforcement surrounding the tool, resulting in a lamination assembly comprising the preform and the tool (204); heating (206) the tool and the preform; co-bonding (208) the tool to the fiber reinforcement within the lamination assembly; solidifying (210) the preform and the tool to form a hybrid composite part; A medium comprising:

[0019] 9. The method comprises: forming the tool such that a body of the tool includes a void; 9. The medium of claim 8, further comprising:

[0020] 10. Fabricating said tooling from a sheet of metal; 10. The medium of any of claims 8 or 9, further comprising:

[0021] 11. Fabricating the tool via additive manufacturing; The medium described in any one of appendixes 8 to 10, further comprising:

[0022] 12. transferring stresses during service of the hybrid composite component from a fibre reinforced portion of the hybrid composite component to the tool; 12. The medium of any one of claims 8 to 11, further comprising:

[0023] 13. Elastically deforming the hybrid composite component in response to an applied stress while maintaining a bond between the tool and a fiber reinforced portion of the hybrid composite component; 13. The medium of any one of claims 8 to 12, further comprising:

[0024] 14. A portion of an aircraft assembled in accordance with a method defined by instructions stored on a computer-readable medium according to any of claims 8 to 13.

[0025] According to further aspects of the present article, there is provided the following:

[0026] 15. A product in the form of a hybrid composite part, Tools (100) and a fiber reinforcement (350) surrounding and co-bonded with said tool; Equipped with The tool encloses a void (150).

[0027] 16. The product of claim 15, wherein the tool forms a core of the hybrid composite part.

[0028] 17. The product of claim 15 or 16, wherein the tool includes notches (160) distributed along its length.

[0029] 18. The product of any one of claims 15 to 17, wherein the notch increases the amount of deflection of the tool.

[0030] 19. The article of manufacture of any one of claims 15 to 18, wherein the tool is made from a material having an elastic modulus between 0.5 and 1.5 times the elastic modulus of the fiber reinforcement.

[0031] 20. The product of any one of claims 15 to 19, wherein the fiber reinforcement is carbon fiber reinforced polymer (CFRP).

[0032] 21. The product of any one of claims 15 to 20, wherein the tool is made from a material lighter than carbon fiber reinforced polymer (CFRP).

[0033] 22. The article of any one of claims 15 to 21, wherein the tool maintains rigidity at the cure temperature of the fiber reinforcement.

[0034] 23. Manufacture of aircraft parts using any of the products described in any of Annexes 15 to 22.

[0035] According to further aspects of the apparatus, there is provided:

[0036] 24. A body (110) having surfaces (120, 130, 140) formed from a core material to a desired dimension and made from a material that maintains its rigidity at the curing temperature of the fiber reinforcement; one or more flanges (112, 114, 116) disposed at the locations where the surfaces are joined together; a void (150) within the body running along the axial length of the body; The apparatus in the form of a hybrid composite part includes tooling (100).

[0037] 25. The apparatus of claim 24, further comprising a notch (160) disposed on each side of the body to reduce the stiffness of the body.

[0038] 26. The apparatus of claim 24 or 25, wherein each face of the body follows a contour (122, 132, 142).

[0039] 27. The apparatus of any one of clauses 24 to 26, wherein the flange is tapered.

[0040] Other embodiments (e.g., methods and computer-readable media related to the above embodiments) are described below. The described features, functions, and advantages can be achieved independently in various embodiments or may be combined in yet other embodiments, further details of which can be understood by referring to the following description and drawings. [Brief description of the drawings]

[0041] Some embodiments of the present disclosure will now be described, by way of example only, with reference to the accompanying drawings, in which like reference numbers represent the same or similar elements in all drawings, and in which:

[0042] [Figure 1] 1 illustrates tooling for a hybrid composite part in an embodiment. [Diagram 2] 1 is a flow chart illustrating a method of utilizing tooling during manufacturing of a hybrid composite part in an embodiment. [Diagram 3] FIG. 1 is a perspective view illustrating a layup of a preform that conforms to a contour defined by internal tooling in an embodiment. [Figure 4] FIG. 2 is an exploded view of tooling made from multiple sheets of core material in an embodiment. [Figure 5-12] 1 illustrates further tooling and preforms in an embodiment. [Figure 13] FIG. 1 is a block diagram of a hybrid composite part according to an embodiment. [Figure 14] FIG. 1 is a flow diagram of an aircraft manufacturing and service method according to an embodiment. [Figure 15] FIG. 1 is a block diagram of an aircraft according to an embodiment. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0043] The drawings and the following description provide specific embodiments of the present disclosure. Thus, it will be understood that those skilled in the art can devise various configurations that embody the principles of the present disclosure and are included within the scope of the present disclosure, although not explicitly described or illustrated herein. Furthermore, the examples described herein are intended to help understand the principles of the present disclosure, and should not be construed as being limited to the examples and conditions specifically described as such. As a result, the present disclosure is not limited to the specific embodiments or examples described below, but is limited by the claims and their equivalents.

[0044] Composite parts, such as carbon fiber reinforced polymer (CFRP) parts, are first laid up in multiple layers, which together are called a preform. The individual fibers within each layer of the preform are aligned parallel to one another, but different layers may exhibit different fiber orientations to enhance strength along different dimensions of the resulting composite. The preform may contain a viscous resin that solidifies to solidify the preform into a composite part (e.g., for an aircraft). Carbon fiber impregnated with uncured thermosetting or thermoplastic resins is called a "prepreg." Another type of carbon fiber is "dry fiber," which is not impregnated with a thermosetting resin but may contain a tackifier or binder. Dry fiber may be infused with resin before curing. For thermosetting resins, solidification is a unidirectional process called curing, while for thermoplastic resins, the resin may reach a viscous state if it is reheated.

[0045] 1 illustrates tooling 100 for a hybrid composite part in an embodiment. As used herein, a "hybrid composite part" is a part that includes a composite material (e.g., a fiber-reinforced composite material such as CFRP) surrounding a core made from a different material.

[0046] Tooling 100 defines a shape for the fiber reinforcement preform and comprises a suitable rigid material capable of withstanding the heat and temperature applied during solidification of the fiber reinforcement (e.g., cure temperatures of approximately 500 degrees Fahrenheit (F) for thermoset parts and approximately 900 degrees Fahrenheit for thermoplastic parts). Tooling 100 may also be referred to herein as a "tool." Tooling 100 is integrated with the fiber reinforcement during solidification. For example, tooling 100 may be made from a sheet of core material, such as metal (e.g., titanium, aluminum), additive manufacturing material added via an additive manufacturing process, etc. In further embodiments, the core material may be lighter in volume weight than solid CFRP. This lighter weight provides technical advantages by reducing weight, which, if the resulting composite part is used in a vehicle, results in reduced fuel costs.

[0047] In yet another embodiment, it may be desirable for the modulus of elasticity and / or coefficient of thermal expansion of the core material to substantially match that of the remainder of the composite part (e.g., to prevent internal stresses in the finished cured part resulting from modulus mismatch). In such an embodiment, the tooling 100 may have a modulus of elasticity between 0.5 and 1.5 times the modulus of elasticity of the fiber reinforcement. In this way, when the tooling 100 forms part of a hybrid composite part, elastic deflection of the hybrid composite part does not damage or break the bond between the tooling 100 and the fiber reinforced portion of the hybrid composite part.

[0048] Tooling 100 includes a body 110 having faces (i.e., faces 120, 130, and 140). Body 110 is elongated in this embodiment, but in further embodiments, body 110 may be formed according to any suitable shape. Each face of tooling 100 is shaped according to a contour. In this embodiment, face 120 is shaped according to contour 122 (which is arcuate), face 130 is shaped according to contour 132 (which is arcuate), and face 140 is shaped according to contour 142 (which is flat). These contours define the shape that a preform placed (e.g., placed or laid up) on tooling 100 will take. Conceptualized in another way, the faces of tooling 100 are complementary to the faces of cavities present in a laminate that is placed on tooling 100. Faces 130 and 120 are rounded due to bends present in the laminate, while face 140 is flat.

[0049] In this embodiment, the notches 160 are distributed along the length of the body 110, resulting in flanges 112, 114, and 116 located where the surfaces are joined. The notches 160 increase the amount of flexure of the tooling 100, allowing it to bend more easily along the longitudinal axis L. The notches increase flexibility, in particular by reducing the amount of material at certain locations. This reduces the bending stiffness / rigidity at these locations, and therefore reduces resistance to bending at these locations. The notches may extend, for example, from half the height to the full height of the flanges, may be several millimeters wide, and may be spaced every few centimeters or meters depending on the amount of flexing desired. For linear tooling 100 that is not subject to dynamic bending loads, notches are not required. Although referred to as "notches," the notches 160 need not be physically cut into the body 110, but rather may include gaps where no material is placed during the fabrication of the body 110.

[0050] This can be advantageous in environments where tooling 100 is subject to flexing after being incorporated into a hybrid composite part. This flexing also allows tooling 100 to flex to accommodate joggles, bends, or twists into the laminate cavities. Tooling 100 also includes voids 150 in this embodiment. Voids 150 are spaces through which cabling or electromagnetic signal transmission lines can pass. For example, voids 150 can facilitate non-destructive testing of tooling 100 via ultrasonics.

[0051] Exemplary details of the operation of tooling 100 are described with respect to Figure 2. In this embodiment, it is assumed that tooling 100 has been fabricated via a mold or via an additive manufacturing process.

[0052] 2 is a flow chart illustrating a method 200 for utilizing tooling during the manufacture of a hybrid composite part, in an embodiment. Although the steps of method 200 are described with reference to tooling 100 of FIG. 1, one of ordinary skill in the art will appreciate that method 200 may also be performed for other tooling that is incorporated into a hybrid composite part. The steps of the flow chart described herein are not all-inclusive and other steps not shown may be included. The steps described herein may also be performed in an alternative order.

[0053] Tooling 100 having a body and surfaces fabricated from a core material to a desired geometry is selected at step 202. Tooling 100 may be selected based on, for example, its shape and the desired shape of the hybrid composite part to be manufactured according to a predetermined schedule, etc.

[0054] In step 204, a fiber reinforcement preform (e.g., preform 350 in FIG. 3) is placed such that it surrounds body 110 (e.g., surrounds the face of tooling 100). This results in a sheet assembly 300 including preform 350 and tooling 100. For example, as shown in sheet assembly 300 in FIG. 3, preform 350 includes sheets 360, 370, and 380 of CFRP that surround the periphery of tooling 100. That is, preform 350 is placed (e.g., laid up, laid flat elsewhere, and then installed, etc.) on each of faces 120, 130, and 140 of tooling 100. Because tooling 100 is rigid, it defines the contours of the sheet of preform 350. Tooling 100 also fills and is stressed through inner corners 352 of preform 350. This enhances the structural strength of the resulting hybrid composite part by enhancing the ability of the part to resist delamination when subjected to stress. For example, as tooling 100 tapers within preform 350, it exhibits an increased amount of interfacial contact between tooling 100 and preform 350. Thus, the taper (e.g., between a ratio of 5:1 and 100:1) provides a technological hardening in the form of increased bond strength based on the increased amount of interfacial contact between tooling 100 and preform 350. The increased amount of interfacial contact increases the area along which tooling 100 and preform 350 are bonded, which in turn increases the bond strength between these parts. While in FIG. 3, portions of tooling 100 are shown as protruding outwardly from preform 350, it will be understood that such portions are merely for illustrative and contextual purposes and that portions of tooling 100 need not protrude outwardly from preform 350.

[0055] In one embodiment, the tooling described herein is also configured to provide support against crush loads as it increases the amount of material in the hybrid composite part that withstands the crush force. The tooling described herein may also provide limited structural support for tensile bending loads in the hybrid composite part.

[0056] With preform 350 placed on tooling 100, such as by placing tooling 100 on sheet 380 or bending sheets 360 and 370 onto tooling 100, preform 350 and tooling 100 may be placed in a vacuum bag for consolidation and curing, or may be placed in a mandrel for consolidation and curing. In step 206, tooling 100 and preform 350 are heated. This may initiate a curing or other solidification process of the resin in preform 350.

[0057] In step 208, the tooling 100 is co-bonded to the fiber reinforcement in the laminar assembly 300. This may result from the solidification process of step 206.

[0058] In step 210, the preform 350 and tooling 100 are solidified to form a hybrid composite part that includes the tooling 100 as an integral component. In a thermosetting environment, the heating in step 206 hardens the resin in the preform 350. This heating operation therefore co-bonds the preform 350 to the tooling 100; that is, the resin in the tooling 100 solidifies and bonds to the core material that forms the tooling 100. In a thermoplastic environment, cooling the preform 350 (after it reaches a molten state) causes the preform 350 to solidify and solidify around the tooling 100. Because the tooling 100 is rigid, the preform 350 does not collapse during the consolidation and / or solidification process. If the tooling 100 was not present or was not rigid, there would be a much greater potential for the preform 350, and any cavities defined by the preform 350, to collapse.

[0059] The method 200 provides technical advantages by resulting in a tooling 100 that defines the shape of the composite part and enhances its strength. For example, the tooling 100 may enhance the strength of the inner corners of the resulting hybrid composite part. This means that the stresses experienced while carrying a load on the hybrid composite part are transferred from the outside of the hybrid composite part (i.e., the fiber reinforced portion) to the tooling that forms the core of the hybrid composite part. Depending on the dimensions and material properties of the hybrid composite part, this facilitates the hybrid composite part to elastically deform in a predetermined and desired manner in response to the applied stresses. Furthermore, since the tooling 100 is incorporated within the preform, the tooling 100 does not need to be removed from the preform after solidification of the preform, which can be a difficult process. This eliminates the labor associated with removing the tooling and prevents the composite part from being damaged by the removal process.

[0060] Figure 3, discussed above with respect to the method of Figure 2, illustrates a blade stiffener or inverted "T" shaped sheet metal assembly 300. However, Figure 3 illustrates only one of many potential versions of a blade stiffener that may be utilized in accordance with the embodiments described herein.

[0061] FIG. 4 is an exploded view of tooling 400 made from multiple sheets of core material in an embodiment. Tooling 400 may be utilized as a mandrel on which preform 350 is placed and solidified. Tooling 400 may thus be utilized in a manner similar to tooling 100. The sheets of core material are welded or bonded together to form tooling 400, and then notches may be added to tooling 400 at a later stage. In this embodiment, sheet 420 is shaped according to contour 122, sheet 430 is shaped according to contour 132, and sheet 440 is shaped according to contour 142. These contours are complementary to the bend radii found in sheets 360 and 370 of preform 350. The manufacturing techniques for forming tooling 100 discussed in FIG. 4 may be used as an alternative to additive manufacturing techniques, if desired.

[0062] 5-12 show additional tooling that is combined with the preform in the embodiment. FIG. 5 shows tooling 500 in the embodiment. In FIG. 5, tooling 500 includes surface 520 that forms a top surface according to contour 522. Tooling 500 also includes surface 540 that forms a bottom surface according to contour 542. Contour 522 defines rounded portion 550 that encloses void 552. Notches 562 and 564 subdivide perimeter 510 of tooling 500 into flange 512. Tooling 500 may enhance the structural strength of composite parts having complex shapes. Alternatively, tooling 500 may be utilized to define passages for cable routing or electromagnetic signal transmission in the resulting hybrid composite part. In one embodiment, void 552 is used for this purpose. Tooling 500 may be utilized to support a laminate that is solidified to form a blade stiffener. One sheet of laminate may be bent around surface 520, while another sheet of laminate may be placed beneath surface 540. Tooling 500 may support the sheets during curing, impart a desired shape to the laminate, and allow one or more voids, such as void 552 and void 554, to be formed as desired.

[0063] Figures 6-7 show tooling for hybrid composite parts having significant vertical dimensions, Figures 8-9 show tooling for hybrid composite parts having triangular cross sections, and Figures 10-12 show tooling for reinforcing hybrid composite parts having D-shaped cross sections.

[0064] 6-7 show a similar blade stiffener configuration as shown in FIG. 3, but with a greater vertical projection of the tooling. This allows for increased support / strengthening of the vertical portions of the blade stringer by increasing the amount of interfacial contact along the vertical portions of the blade stringer between the laminate and the tooling. Thus, the interfacial strength of the bond between the tooling and the laminate is increased, which in turn increases the resistance to delamination of the resulting hybrid composite part.

[0065] In FIG. 6, tooling 600 includes vertical extensions that increase the strength of the resulting hybrid composite part. Tooling 600 includes body 610 having face 620 (following contour 622), face 630 (following contour 632), and face 640 (following contour 642). Tooling 600 surrounds void 650 and includes flanges 612, 614, and 616 formed by notches 662 and 664. In FIG. 7, tooling 600 provides layup 700 with a shape for preform 750. Tooling 600 also increases the structural strength of the inner corners 752 of preform 750 because it supports these inner corners 752. This means that stresses that could pull sheet 760 away from sheet 780 are transferred through tooling 600 and not through inner corners 752. Because tooling 600 is bonded to the sheets along a significant distance, tooling 600 has significant interfacial bond strength with the sheets. At the same time, because the radius of the inner corners of tooling 600 is greater than the radius of inner corners 752 of preform 750, and because the inner corners of tooling 600 are formed from a single piece rather than from two separate sheets bonded together, the stresses experienced at the inner corners of tooling 600 are more likely to be sustained. Inner corners 752 are formed by the intersections between sheets 760, 770, and 780 of preform 750. Tooling 600 may also provide a geometric structural advantage by increasing the distance from center 790 of the hybrid composite part to inner corners 752. This increased distance reduces the peel / separation forces experienced by the hybrid composite part.

[0066] In FIG. 8, tooling 800 defines a large triangular cavity (i.e., void 850). Tooling 800 includes a body 810 having face 820 (following contour 822), face 830 (following contour 832), and face 840 (following contour 842). Tooling 800 encloses void 850 formed by notch 862 and notch 864 and has flanges 812, 814, and 816. In FIG. 9, tooling 800 imparts a shape for preform 950 to layup 900. Tooling 800 also enhances the structural strength of inner corner 952 of preform 950, similar to tooling 600 of FIG. 6 described above. Inner corner 952 is formed by the intersection between sheets 960, 970, and 980 of preform 950.

[0067] In FIG. 10, tooling 1000 is formed according to the shape of a D. In this embodiment, tooling 1000 includes a body 1010 having a surface 1020 (following contour 1022) and a surface 1040 (following contour 1042). Tooling 1000 surrounds a void 1050 and has flanges 1012, 614, and 616 formed by notches 1062. In FIG. 11, tooling 1000 imparts a shape for preform 1150 to layup 1100. Tooling 1000 also enhances the structural strength of inner corner 1152 of preform 1150, similar to tooling 600 of FIG. 6 described above. Inner corner 1152 is formed by the intersection of sheets 1160 and 1170 of preform 1150.

[0068] Figure 12 illustrates a further embodiment in which tooling 1000 may be utilized. In Figure 12, a first set 1210 of tooling 1000 located on the right side of layup 1200 supports a series of D-shaped portions of preforms 1250, while a second set 1220 of tooling 1000 is inverted relative to the first set 1210 and supports another series of D-shaped portions of preforms 1250. EXAMPLES

[0069] The following examples describe additional processes, systems, and methods in the context of hybrid composite parts.

[0070] FIG. 13 is a block diagram of a hybrid composite part 1300 in one embodiment. According to FIG. 13, the hybrid composite part 1300 includes a tooling 1330 made from a material 1360, such as titanium, and enclosing a void 1370. A preform 1350 is placed on a surface 1332 of the tooling 1330. The preform 1350 includes a ply 1310 (including resin 1312 and fibers 1314). The preform 1350 also includes a ply 1320 (including resin 1322 and fibers 1324). The tooling 1330 provides a rigid structure onto which the preform 1350 may be placed and solidified. Additionally, the tooling 1330 increases the strength of the inner corners of the preform 1350, for example, by increasing the radius of the inner corners of the preform 1350. That is, radius 1334 of tooling 1330 is greater than the radius formed by inside corner 1352 at the intersection between ply 1320 and ply 1310 .

[0071] With more particular reference to the drawings, embodiments of the disclosure will be described in the context of aircraft manufacturing and service in method 1400 as shown in FIG. 14 and aircraft 1402 as shown in FIG. 15. Before production begins, method 1400 includes specification and design 1404 of the aircraft 1402, and material procurement 1406. During production, component and subassembly manufacturing 1408 and system integration 1410 of the aircraft 1402 occurs. The aircraft 1402 then undergoes certification and delivery 1412 and enters service 1414. While in service with a customer, the aircraft 1402 is scheduled for periodic maintenance and service 1416 activities, which may also include modifications, reconfigurations, refurbishments, etc. Apparatus and methods embodied herein may be employed during one or more appropriate stages of production and service described in method 1400 (e.g., specification and design 1404, material procurement 1406, component and subassembly manufacturing 1408, systems integration 1410, certification and delivery 1412, in-service 1414, maintenance and service 1416) and / or in any appropriate component of aircraft 1402 (e.g., airframe 1418, systems 1420, interiors 1422, propulsion system 1424, electrical system 1426, hydraulic system 1428, environmental system 1430).

[0072] Each process of method 1400 may be performed or implemented by a system integrator, a third party, and / or an operator (e.g., a customer). For purposes of this description, a system integrator may include, but is not limited to, any number of aircraft manufacturers and subcontractors of major systems, a third party may include, but is not limited to, any number of vendors, subcontractors, and suppliers, and an operator may be an airline, a leasing company, a military entity, a service organization, etc.

[0073] 15 , an aircraft 1402 produced by method 1400 may include an airframe 1418 with a number of systems 1420 and an interior 1422. Examples of systems 1420 include one or more of a propulsion system 1424, an electrical system 1426, a hydraulic system 1428, and an environmental system 1430. Any number of other systems may be included. Although an aerospace example is provided, the principles of the invention may be applied to other industries, such as the automotive industry.

[0074] As previously mentioned, apparatus and methods embodied herein may be used during any one or more of the stages of manufacturing and in-service described in method 1400. For example, components and subassemblies corresponding to component and subassembly manufacturing 1408 may be fabricated or manufactured in a manner similar to components or subassemblies manufactured during the in-service life of aircraft 1402. Also, one or more apparatus embodiments, method embodiments, or combinations thereof may be utilized during subassembly manufacturing 1408 and during system integration 1410, such as by substantially speeding up or reducing the cost of assembly of aircraft 1402. Similarly, one or more apparatus embodiments, method embodiments, or combinations thereof may be utilized during the in-service life of aircraft 1402, such as, but not limited to, during maintenance and service 1416. For example, the techniques and systems described herein may be used for material procurement 1406, component and subassembly manufacturing 1408, systems integration 1410, in-service 1414, and / or maintenance and service 1416, and / or may be used for the airframe 1418 and / or interiors 1422. These techniques and systems may even be utilized for systems 1420, including, for example, propulsion system 1424, electrical system 1426, hydraulic system 1428, and / or environmental system 1430.

[0075] In one embodiment, the parts include a portion of the airframe 1418 and are manufactured during part and subassembly manufacturing 1408. The parts are then assembled into the aircraft during system integration 1410 and may then be utilized during in-service 1414 until wear renders the parts unusable. The parts are then scrapped and replaced with newly manufactured parts during maintenance and service 1416. The parts and methods of the present invention may be utilized throughout part and subassembly manufacturing 1408 to manufacture new parts.

[0076] Any of the various control elements (e.g., electrical or electronic components) shown or described herein may be implemented as hardware, software-implemented processors, firmware-implemented processors, or any combination thereof. For example, an element may be implemented as dedicated hardware. A dedicated hardware element may be referred to as a "processor," "controller," or other similar terms. If implemented by a processor, the functionality may be provided by a single dedicated processor, or by a single shared processor, or by multiple separate processors, some of which may be shared. Furthermore, explicit use of the terms "processor" or "controller" should not be construed to refer only to hardware capable of executing software, but may implicitly include, without limitation, digital signal processor (DSP) hardware, network processors, application specific integrated circuits (ASICs) or other circuitry, field programmable gate arrays (FPGAs), read only memory (ROM) for storing software, random access memory (RAM), non-volatile storage, logic circuits, or other physical hardware components or modules.

[0077] Additionally, a control element may be implemented as instructions executable by a processor or computer to perform the functions of the element. Some examples of instructions include software, program code, and firmware. These instructions operate when executed by a processor to direct the processor to perform the functions of the element. The instructions may be stored in a storage device readable by the processor. Examples of storage devices include digital or solid-state memory, magnetic storage media such as magnetic disks or magnetic tapes, hard drives, or optically readable digital data storage media.

[0078] Although specific embodiments have been described herein, the scope of the disclosure is not limited to these specific embodiments, and is defined by the following claims and their equivalents.

Claims

1. selecting a tool having a flange defining a surface made from a core material with a desired geometry; placing a fiber reinforced preform having an internal corner around the tool such that the flange of the tool fills the internal corner of the preform, resulting in a lamination assembly including the preform and the tool; heating the tool and the preform; co-bonding the tool to the fiber reinforcement within the lamination assembly; consolidating the preform and the tool to form a hybrid composite part; A method for manufacturing a composite part, comprising:

2. The method of claim 1 further comprising forming the tool such that a body of the tool includes a void.

3. transferring stresses during service of the hybrid composite component from the fiber reinforcement of the hybrid composite component to the tool; The method of claim 1 or 2, further comprising:

4. 4. The method of claim 1, further comprising the step of elastically deforming the hybrid composite component in response to an applied stress while maintaining a bond between the tool and a fiber reinforced portion of the hybrid composite component.

5. A non-transitory computer readable medium embodying program instructions operable, when executed by a processor, to perform a method of manufacturing a composite part, the method comprising: selecting a tool having a flange defining a surface fabricated from a core material to a desired dimension; placing a fiber reinforced preform having an internal corner around the tool such that the flange of the tool fills the internal corner of the preform, resulting in a lamination assembly comprising the preform and the tool; heating the tool and the preform; co-bonding the tool to the fiber reinforcement within the lamination assembly; solidifying the preform and the tool to form a hybrid composite part; Consisting of Medium.

6. Tools and a fiber reinforcement surrounding the tool and co-bonded within the tool; Including, the tool enclosing a gap; The fiber reinforcement has an inner corner, the tool has a plurality of flanges that fill the interior corners of the fiber reinforcement; A product in the form of a hybrid composite part, wherein each flange has a plurality of notches distributed along the length of the tool to increase the amount of deflection of the tool, and each notch in each flange extends transversely of the tool toward the other flange.

7. The article of manufacture of claim 6 , wherein the tool forms a core of the hybrid composite part.

8. The tool has a modulus of elasticity between 0.5 and 1.5 times the modulus of elasticity of the fiber reinforcement.

8. An article according to claim 6 or 7 made from a material.

9. 9. An article according to any one of claims 6 to 8, wherein the tool maintains its rigidity at the cure temperature of the fibre reinforcement.

10. Manufacture of a part of an aircraft using a product according to any one of claims 6 to 9.

11. a fiber reinforcement having an inner corner; a tool surrounded by the fiber reinforcement, the tool comprising: a body made of a material having a surface formed to a desired dimension from a core material and maintaining rigidity at a curing temperature of the fiber reinforcement; a plurality of flanges disposed at positions where the surfaces are joined together and filling the inner corners of the fiber reinforcement; a void within the body running along an axial length of the body; each flange has a plurality of notches distributed along the length of the tool to increase the amount of deflection of the tool, each notch in each flange extending transversely of the tool toward the other flange; The device in the form of a hybrid composite part.

12. The apparatus of claim 11 , wherein the flange is tapered.

Images