Composite spar with integrated sacrificial surface

Abstract

To provide composite spars with integrated sacrificial surfaces.SOLUTION: Composite assemblies are described that include composite spars that are co-cured with one or more sacrificial members on their flanges, forming an integrated sacrificial surface for the composite spars. In one embodiment, the composite assembly includes a composite spar having a web and flanges that project from sides of the web. The composite assembly further includes a sacrificial member of composite materials co-cured with the composite spar on an outer surface of at least one of the flanges. In addition, the sacrificial member has an outer surface that has been machined into conformance with an inner surface of at least one skin panel for an aircraft structure to form a contact surface with the at least one skin panel.SELECTED DRAWING: Figure 2

Description

[Technical Field]

[0001] The present disclosure relates to the field of aircraft structures, and more particularly to aircraft structures utilizing spars as structural members. [Background technology]

[0002] In the assembly process of an aircraft wing, ribs and spars are assembled to form the skeleton of the wing. The spars generally form the main structural members of the wing and extend along the length of the wing. The ribs are attached to the spars (e.g., a rib may be attached to the front spar at the leading edge of the wing and the rear spar at the trailing edge of the wing) and generally have an outer periphery shape that defines the wing's airfoil. The outer surface of the wing is formed by skin panels that are attached along the periphery of the ribs and the periphery of the spars to form a smooth surface for the wing.

[0003] To ensure proper fit of the skin panel to the spar, shims can be installed around the perimeter of the spar where the inner surface of the skin panel meets the spar. The shims are used to eliminate possible machining processes on the spar itself that could adversely affect the structural integrity of the spar.

[0004] Typically, installing shims into spars is a time-consuming manual process that requires additional man-hours to fit the shims into the gaps between the skin panels and the spars, adding to the assembly and disassembly time of the wing or other aircraft structure that uses the spars. Summary of the Invention [Problem to be solved by the invention]

[0005] Therefore, based on the foregoing, it remains desirable to improve aircraft manufacturing processes, and in particular to improve processes for assembling aircraft structures including spars. [Means for solving the problem]

[0006] A composite assembly is described that includes a composite spar that is co-cured with one or more sacrificial members on a flange to form an integrated sacrificial surface for the composite spar. Generally, a spar is an elongated structure that includes a web that defines the spar's main surface and flanges that protrude from the sides of the web. The spar extends spanwise within a wing at a right angle or nearly right angle to the fuselage and forms the wing's primary structural component. The web within the spar is positioned vertically within the wing, and the flanges define surfaces for attaching skin panels. During assembly, gaps may exist between the flanges and the skin panels. In the embodiments described herein, the sacrificial members co-cured with the flanges can be machined to match the outer surface of the sacrificial member with the surface of the skin panels, providing the technical advantage of reducing gaps between the spar and the skin panels within a wing or other aircraft structure utilizing a spar.

[0007] One embodiment includes a composite assembly including a composite spar having a web and a flange projecting from a side of the web. The composite assembly further includes a sacrificial member of composite material co-cured with the composite spar on at least one outer surface of the flange. The sacrificial member further includes an outer surface machined to match an inner surface of at least one skin panel for an aircraft structure to form an interface with the at least one skin panel.

[0008] Another embodiment includes a method of manufacturing a composite assembly. The method includes assembling a first composite layup to define web and flange portions for a composite spar of the composite assembly, assembling a second composite layup over at least one of the flange portions to define a sacrificial member for the composite assembly, and co-curing the first and second composite layups to solidify the composite assembly. The method further includes machining an outer surface of the sacrificial member to match an inner surface of at least one skin panel for the aircraft structure to form a contact surface for the at least one skin panel.

[0009] Another embodiment includes a method of manufacturing a composite assembly. The method includes performing a first composite layup on a layup mandrel defining a contour of a composite spar of the composite assembly, performing a second composite layup on a flange portion of the first composite layup defined by the contour, and co-curing the first composite layup and the second composite layup to solidify the composite assembly, wherein the first composite layup forms the composite spar and the second composite layup forms a sacrificial member on the flange on the composite spar. The method further includes calculating a machining depth of the sacrificial member based on an estimated clearance tolerance between the flange of the composite spar and at least one skin panel for an aircraft structure, and machining an outer surface of the sacrificial member along at least a portion of the length of the composite spar based on the machining depth to form a contact surface of the at least one skin panel.

[0010] The features, functions, and advantages described above can be achieved independently in various embodiments or can be combined in yet other embodiments, further details of which can be seen with reference to the following description and drawings.

[0011] Some embodiments will now be described, by way of example only, with reference to the accompanying drawings, in which the same reference numbers represent the same elements or types of elements in all drawings, and in which: [Brief explanation of the drawings]

[0012] [Figure 1] 1 shows an aircraft wing according to the prior art; [Figure 2] FIG. 1 illustrates an isometric view of a composite assembly in an exemplary embodiment. [Figure 3] 3 is a cross-sectional view of the composite assembly taken along section line AA of FIG. 2 in an exemplary embodiment. [Figure 4] 1 illustrates a portion of an aircraft structure in an illustrative embodiment; [Figure 5] 5 is a diagram of the region of FIG. 4 in an exemplary embodiment. [Figure 6] 1 is a method of manufacturing a composite structure in an exemplary embodiment. [Figure 7] 7 illustrates further details of the method of FIG. 6 in an exemplary embodiment. [Figure 8] 7 illustrates further details of the method of FIG. 6 in an exemplary embodiment. [Figure 9] 7 illustrates further details of the method of FIG. 6 in an exemplary embodiment. [Figure 10] 7 illustrates further details of the method of FIG. 6 in an exemplary embodiment. [Figure 11] 7 illustrates further details of the method of FIG. 6 in an exemplary embodiment. [Figure 12] 7 illustrates further details of the method of FIG. 6 in an exemplary embodiment. [Figure 13] 7A-7C are isometric views of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 14] 7A-7C are isometric views of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 15] 7A-7C are isometric views of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 16] 7A-7C are isometric views of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 17] 7A-7C are isometric views of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 18] 7A-7C are isometric views of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 19] 7A-7C are isometric views of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 20] 7A-7C are isometric views of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 21] 7A-7C are isometric views of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 22] 7A-7C are isometric views of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 23]7A-7C are isometric views of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 24] 7A-7C are isometric views of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 25] 10 is a flowchart of another method for manufacturing a composite structure in an illustrative embodiment. [Figure 26] 26 illustrates further details of the method of FIG. 25 in an exemplary embodiment. [Figure 27] 3 is a cross-sectional view of the composite assembly taken along section line CC of FIG. 2 in an exemplary embodiment. [Figure 28] 3 is a cross-sectional view of the composite assembly taken along section line CC of FIG. 2 in an exemplary embodiment. [Figure 29] 1 is a flowchart illustrating an aircraft manufacturing and service method in an illustrative embodiment. [Figure 30] 1 is a diagrammatical illustration of an aircraft in an illustrative embodiment; DETAILED DESCRIPTION OF THE INVENTION

[0013] The drawings and the following description illustrate certain exemplary embodiments. Those skilled in the art will understand that, although not explicitly described or shown herein, they can devise various configurations that embody the principles described herein and fall within the scope of the contemplated claims that follow this description. Furthermore, any examples described herein are intended to aid in the understanding of the principles of the present disclosure and should not be construed as limiting. Consequently, the present disclosure is not limited to the specific embodiments or examples described below, but rather by the claims and their equivalents.

[0014] FIG. 1 illustrates a prior art wing 100. The wing 100 includes a plurality of ribs 102 extending between a front spar 104 and an aft spar 105, with the front spar 104 near a leading edge 106 of the wing 100 and the aft spar 105 toward a trailing edge 107 of the wing 100. In FIG. 1, the ribs 102 are spaced along the length of the wing 100 and their periphery forms the airfoil shape of the wing 100. In FIG. 1, the wing 100 includes a nose rib 108 extending from the front spar 104 toward the leading edge 106 of the wing 100 to form the front shape of the airfoil at the leading edge 106. The wing 100 also includes a skin 110, which is partially removed in this view to show the structure underlying the wing 100.

[0015] As is apparent from FIG. 1, the ribs 102 and spars 104-105 form the framework of the wing 100, and the skin 110 takes the shape of the ribs 102 along the peripheral surface between the spars 104-105.

[0016] 2 is an isometric view of an exemplary embodiment of a composite assembly 200. In this embodiment, composite assembly 200 includes a composite spar 202 and one or more sacrificial members 204 of composite material that are co-cured to form a structural component for an aircraft, such as an aircraft wing, an aircraft empennage, or an aircraft tail section.

[0017] Composite parts, such as carbon fiber reinforced polymer (CFRP) and / or glass fiber reinforced polymer (GFRP) parts, are initially laid up in multiple layers of material, sometimes called reinforcing layers. The individual fibers within each layer are aligned parallel to one another, but different layers exhibit different fiber orientations along different dimensions to increase the strength of the resulting composite part. The layers are pre-impregnated with a matrix material, such as an uncured thermoset or thermoplastic resin, and are called "prepregs." Alternatively, the layers may be laid up dry (i.e., "dry fiber") and then infused with the matrix material before curing.

[0018] Referring again to Figure 2, composite spar 202 in this embodiment is an elongated member that extends generally in the shape of an inverted "U" along length 206. Composite spar 202 is substantially planar and includes web 208 that defines a major surface 210 of composite spar 202. In this embodiment, flanges 214-215 project downwardly from sides 212-213 of web 208 in Figure 2 and define a width 216 of composite spar 202 and a height 218 of composite spar 202.

[0019] In this embodiment, sacrificial member 204 is co-cured to flange 214 and / or flange 215. After composite assembly 200 is cured, outer surface 220 of sacrificial member 204 can be machined as needed during the manufacturing process to mitigate gaps that may occur between flange 214 and / or flange 215 of composite spar 202 and other aircraft structures, such as skin panels. The use of sacrificial member 204 provides the technical benefit of reducing the time and effort required to assemble various aircraft structures that use spars and skin panels by mitigating the need for manual shims that would normally be introduced when gaps exist when a spar flange is mated to a skin panel.

[0020] In general, it is desirable for the sacrificial member 204 to have a stiffness substantially less than that of the composite spar 202, because removing material from the sacrificial member 204 can minimize changes in the overall stiffness of the composite assembly 200. To achieve this type of result, the composite spar 202 may be formed, for example, from BMS8-276 tape (a CFRP material) using a quasi-isotropic layup, which has a very high stiffness. The sacrificial member 204 may be formed, for example, from BMS8-276 fabric (another CFRP material), which has a moderate stiffness compared to the BMS8-276 tape. The BMS8-276 fabric for the sacrificial member 204 can use various + / - 45 degree orientations during layup to minimize the stiffness of the composite assembly 200 in the spanwise direction. If the sacrificial member 204 is formed from a GFRP having a lower stiffness compared to either the BMS8-276 tape or fabric, various types of materials and orientations of these materials can be realized.

[0021] In this embodiment, sacrificial member 204 is shown on both flanges 214 - 215 , however, in some embodiments, sacrificial member 204 may be formed on either flange 214 or flange 215 .

[0022] Figure 3 is a cross-sectional view of the composite assembly 200 taken along section line AA in Figure 2 in an exemplary embodiment. In this embodiment, the composite spar 202 includes a bend 302 at the side 212 of the web 208 between the web 208 and the flange 214, and the sacrificial member 204 is disposed on the outer surface 304 of the flange 214. The sacrificial member 204 extends from the end 306 of the flange 214 toward the bend 302 and terminates at an edge 308 proximate the bend 302. In this embodiment, the edge 308 is a square edge, although in other embodiments, the edge 308 may have other shapes, such as a taper or a ramp. The use of a taper or ramp on the edge 308 of the sacrificial member 204 can reduce the generation of compression waves in the composite spar 202 during manufacturing, as described below.

[0023] The sacrificial member 204 has an initial thickness 310 defined by the distance between the outer surface 304 of the flange 214 and the outer surface 220 of the sacrificial member 204. Typically, the initial thickness 310 is pre-selected based on expected tolerances between the flange 214 and other components on the aircraft, such as skin panels. After the composite assembly 200 is co-cured, the outer surface 220 of the sacrificial member 204 may be machined (e.g., material is removed from the outer surface 220 of the sacrificial member 204 along the length 206 or a portion of the length 206 of the composite spar 202 (see FIG. 2 )), thereby reducing the initial thickness 310 of the sacrificial member 204 to a final thickness 314 based on a machining profile or machining depth. Machining the sacrificial member 204 also creates an interface 316 with the skin panel. Although layers are not shown in FIG. 3 , both the composite spar 202 and the sacrificial member 204 are formed from composite materials, such as CFRP plies.

[0024] Similar to the flange 214, the composite spar 202 includes a bend 303 between the web 208 and the flange 215 at the side 213 of the web 208, with the sacrificial member 204 disposed on the outer surface 305 of the flange 215. The sacrificial member 204 in this embodiment extends from the end 307 of the flange 215 toward the bend 303 and terminates at an edge 308 proximate the bend 303. In this embodiment, the edge 308 is a square edge, but in other embodiments, the edge 308 may have other shapes, such as a taper or ramp, for reasons similar to those of the edge 308 of the sacrificial member 204 on the flange 214. The sacrificial member 204 on the flange 215 has an initial thickness 310 defined by the distance between the outer surface 305 of the flange 215 and the outer surface 220 of the sacrificial member 204, which may be the same as or different from the initial thickness 310 of the sacrificial member 204 on the flange 214. Typically, initial thickness 310 is selected based on expected tolerances between flange 215 and other components on the aircraft, such as skin panels, and then outer surface 220 of sacrificial member 204 may be machined (e.g., material is removed from outer surface 220 of sacrificial member 204 at flange 215 along length 206 of composite spar 202 (see FIG. 2 )), thereby reducing initial thickness 310 of sacrificial member 204 to final thickness 314 based on the machining profile. Additionally, the amount of material removed from sacrificial member 204 may vary along length 206 or a portion of length 206 of composite spar 202 and / or may vary depending on whether sacrificial member 204 is located on flange 214-215.

[0025] 4 illustrates a portion of an aircraft structure 400 in an exemplary embodiment. In this embodiment, aircraft structure 400 includes composite assembly 200 and skin panels 402-403. In this embodiment, sacrificial member 204 of flange 214 is positioned proximate inner surface 404 of skin panel 402, and sacrificial member 204 of flange 215 is positioned proximate inner surface 405 of skin panel 403 in this embodiment.

[0026] As previously mentioned, shims have often been manually installed between the spar flange and the skin panel during the assembly process to compensate for or mitigate gaps that form between these two components. Generally, direct machining on the composite spar 202 is undesirable because the machining process can remove fiber layers and compromise the structural integrity of the composite spar 202. While the use of shims precludes direct machining on the spar, the use of shims is a time-consuming process that involves temporarily assembling the components together, measuring any gaps that may exist between the spar flange and its skin panel, disassembling the component, and bonding the shim to the spar flange. The shim can then be machined to achieve the final fit between the spar flange and the skin panel. Final assembly of the spar and skin panel can then be performed.

[0027] Once the sacrificial member 204 is co-cured with the flanges 214 and / or 215 of the composite spar 202, a machining process can be performed on the exterior surface 220 of the sacrificial member 204 prior to assembling the aircraft structure 400, thereby saving time and effort over traditional manual shimming processes. For example, the machining profile or machining depth of the sacrificial member 204 can be generated based on several different factors, including the expected tolerances between the flanges 214-215 of the composite spar 202 and the interior surfaces 404-405 of the skin panels 402-403. Once the machining profile or machining depth is selected and the composite assembly 200 is tested for fit against the skin panels 402-403, subsequent spars with integrated sacrificial members can be machined to the same profile or depth, ensuring that each spar is interchangeable between different builds of the same aircraft structure. If gaps are found between the skin panels 402-403 and their corresponding sacrificial members 204 after machining and during assembly, the machining profile or machining depth can be adjusted for subsequent manufacturing of the composite assembly 200 to mitigate the gaps in future builds of the aircraft structure 400.

[0028] In another example, a three-dimensional (3D) scan may be performed on the composite assembly 200 and skin panels 402-403, which may be used to determine the machining profile to apply to the sacrificial member 204. In this case, the composite assemblies 200 and skin panels 402-403 may be serially numbered or marked for use as a group, thereby reducing the likelihood that the composite assemblies 200 are interchangeable between different build instances of the same aircraft structure 400. Both of these different types of profile generation processes are described in more detail below.

[0029] FIG. 5 is a diagram of region 406 of FIG. 4 in an exemplary embodiment. In this embodiment, edge 308 of sacrificial member 204 is in the shape of a ramp or taper proximate bend 303 of composite spar 202. The use of a taper or ramp on edge 308 may be used to mitigate deformation of composite spar 202 that may occur during the manufacturing process, which may cause waves in the fiber plies that make up composite spar 202 due to vacuum compression during curing. Edge 308 may be cut into a ramp shape after layup, or may be laid up as a ramp during manufacturing. When edge 308 is laid up as a ramp, the width of the plies may decrease with increasing distance from outer surface 305 of flange 215 (e.g., by making plies 502 proximate outer surface 305 of flange 215 wider than plies 504 distal to outer surface 305 of flange 215). FIG. 5 also shows the contact surface 316 that contacts the inner surface 405 of the skin panel 403, which is created after machining the outer surface 220 (see FIG. 3) of the sacrificial member 204 until the final thickness 314 is achieved.

[0030] Figure 6 illustrates an exemplary embodiment of a method 600 for manufacturing a composite structure, Figures 7-12 illustrate additional details of the exemplary embodiment of the method 600, and Figures 13-22 illustrate isometric views of various stages of the manufacturing process in an exemplary embodiment. Although the methods described herein are described with respect to various embodiments of the composite assembly 200 and aircraft structure 400, the methods may be applied to other configurations of the composite assembly 200 and aircraft structure 400 not shown or described. The steps of the methods described herein may include other steps not shown, and the steps may be performed in an alternating order.

[0031] Step 602 includes assembling a first composite layup that defines a portion of composite spar 202. In one example, the first composite layup may be formed flat and molded onto a layup mandrel 1302 that defines the outline of composite spar 202 (see FIG. 13 ). Layup mandrel 1302 includes a flat region 1304 that defines the shape of web 208 of composite spar 202, a first side region 1306 that defines the shape and orientation of flange 214 relative to web 208 of composite spar 202, and a second side region 1308 that defines the shape and orientation of flange 215 relative to web 208 of composite spar 202.

[0032] In another example of assembling the first composite layup, an automated fiber placement (AFP) machine may perform the layup directly on the layup mandrel 1302. FIG. 14 shows the first composite layup 1402 on the layup mandrel 1302, defining the web portion 1404, flange portions 1406-1407, and bends 1408-1409 of the composite spar 202 defined by the underlying shape or contour of the layup mandrel 1302. Generally, the first composite layup 1402 includes continuous fiber layers formed as prepregs or dry layups, whose orientation may vary from layer to layer. Furthermore, the layers may be continuous from the end 306 of the flange 214 to the end 307 of the flange 215. As previously mentioned, the first composite layup 1402 may include BMS8-276 tape, which has a quasi-isotropic layup that has very high stiffness after curing.

[0033] Step 604 includes assembling a second composite layup 1502 (see FIG. 15 ) over the flange portion 1406 and / or flange portion 1407 of the first composite layup 1402. For example, a flat layup may be performed on the second composite layup 1502 and then transferred to or directly contacting the flange portion 1406 and / or flange portion 1407 of the first composite layup 1402. In another example, the AFP machine may perform the layup directly over the flange portion 1406 and / or flange portion 1407 of the first composite layup 1402. As previously mentioned, the second composite layup 1502 may include a GFRP material having a medium stiffness after cure compared to BMS8-276 tape and a low stiffness after cure compared to BMS8-276 fabric laminated at varying + / −45 degree orientations, or compared to both BMS8-276 tape or BMS8-276 fabric.

[0034] Step 606 includes co-curing the first composite layup 1402 and the second composite layup 1502 to consolidate the composite assembly 200. For example, the first composite layup 1402 and the second composite layup 1502 may be bagged and placed under vacuum to apply pressure to the first composite layup 1402 and the second composite layup 1502, and heat may be applied to consolidate the composite assembly 200. After curing, the resulting structure is the composite assembly 200 as shown in FIG. 2 , where the composite spar 202 has the web 208 formed from the web portion 1404 and the flanges 214-215 formed from the flange portions 1406-1407, respectively, of the first composite layup 1402, and the sacrificial member 204 is formed from the second composite layup 1502 on one or both of the flanges 214-215.

[0035] Step 608 includes machining the outer surface 220 of the sacrificial member 204 to conform to the surface of one or more skin panels. For example, machining the outer surface 220 of the sacrificial member 204 is performed to remove material from the sacrificial member 204 and form the contact surface 316 of skin panel 402 and / or skin panel 403. Prior to the machining process, the initial thickness 310 of the sacrificial member 204 is Approximately 0.203cm~0.305cm (approximately 0.08"~0.12") After the machining process, the final thickness 314 of the sacrificial member 204 may be Approximately 0.076cm to 0.203cm (approximately 0.03 inches to 0.08 inches) may be.

[0036] In some cases, it may be desirable for the edge 308 of the sacrificial member 204 proximate the bend 302 and / or bend 303 of the composite spar 202 to have a particular shape to prevent structural changes to the composite spar 202 during the manufacturing process of the composite assembly 200. In this case, the edge 1504 of the second composite layup 1502 (see FIG. 15 ) can be formed or cut into a tapered or ramped shape (see step 702 of FIG. 7 ) to align with the bends 1408-1409, respectively, of the first composite layup 1402 (see step 704 of FIG. 7 ). Using a tapered or ramped shape on the edge 1504 of the second composite layup 1502 can prevent waves in the first composite layup 1402 when compression is applied during curing. The ramp or taper can have a particular ratio, such as 1:1 (see step 802 of FIG. 8 ), and the ramp or taper will form a generally 45-degree angle after curing, as shown in FIG. 5 . However, in other embodiments, other types of shapes for the edge 1504 of the second composite layup 1502 may be used to create any type of edge 308 desired for the sacrificial member 204 after curing.

[0037] If manual placement of the second composite layup 1502 onto the first composite layup 1402 is performed before curing, the second composite layup 1502 can be assembled onto an adhesive film 1602 placed on a work surface 1604 (see step 902 in FIG. 9 ), as shown in FIG. 16. As shown in FIG. 16, the second composite layup 1502 includes multiple plies 1606 stacked vertically to achieve a desired thickness 1608 of the second composite layup 1502, which is used to define the initial thickness 310 after curing (see FIG. 3 ). The outer surface 1610 of the second composite layup 1502 forms the outer surface 220 of the sacrificial member 204 after the composite assembly 200 is cured.

[0038] After assembling the second composite layup 1502 onto the adhesive film 1602, the second composite layup 1502 and adhesive film 1602 are transferred to the flange portion 1406 and / or flange portion 1407 of the first composite layup 1402 during the assembly process (see step 904), as shown in FIG. 17, with the adhesive film 1602 contacting the outer surface 304 of the flange 214 and / or the outer surface 305 of the flange 215. Vacuum bagging, compression, and curing can be performed with the adhesive film 1602 disposed between the sacrificial member 204 and the flange 214 and / or flange 215 to produce the composite assembly 200 of FIG. 18. FIG. 19 shows a cross-sectional view of FIG. 18 along section line BB, illustrating the placement of the adhesive film 1602 between the sacrificial member 204 and the flange 214 and / or flange 215.

[0039] 9 and 16, the second composite layup 1502 can be assembled onto the adhesive film 1602, a taper can be cut at the edge 1504 of the second composite layup 1502 (see step 1002 and FIG. 20), and the second composite layup 1502 can then be transferred (see step 904) to the flange portion 1406 and / or flange portion 1407 of the first composite layup 1402 with the edge 1504 aligned with the bend 1408 and / or bend 1409 (see step 1004 and FIG. 17). After curing, the edge 1504 of the second composite layup 1502 forms the edge 308 of the sacrificial member 204 of the composite assembly 200, as shown in the cross section of the composite assembly 200 in FIG. 21.

[0040] As mentioned above, there are numerous processes that can be used to determine how much material to remove from the exterior surface 220 of the sacrificial member 204 after the composite assembly 200 has cured and hardened. One process involves connecting or associating the composite assembly 200 with particular skin panels 402-403 of the aircraft structure 400 (see FIG. 4). First, one or more skin panels to be bonded or mated to the composite spar 202 can be identified (see step 1102 of FIG. 11). Depending on the design of the structure in which the composite spar 202 is to be used, multiple skin panels may be bonded or mated to the composite spar 202 along its length 206. For example, skin panel 403 may be identified as assigned to install on flange 215 of the composite spar 202 (see FIG. 5). The interference between the skin panel 403 and the exterior surface 220 of the sacrificial member 204 can be calculated (see step 1104). For example, a 3D scan of the skin panel 403 and the composite assembly 200 may be performed and used to calculate the interference. The outer surface 220 of the sacrificial member 204 may then be machined based on the interference (see step 1106). For example, the outer surface 220 of the sacrificial member 204 may have an initial pre-machining thickness 310, as shown in FIG. 22, which is based on the thickness 1608 of the uncured second composite layup 1502 (see FIG. 16). The outer surface 220 may then be machined to a final thickness 314, as shown in FIG. 22, before assembling the composite spar 202 and skin panel 403 together. A machining depth 2202, which cuts into the sacrificial member 204 from the outer surface 220 toward the flange 215, removes material from the sacrificial member 204 until the desired final thickness 314 is achieved. While FIG. 22 merely illustrates a cross-section of a particular portion of the aircraft structure 400 for illustrative purposes, the machining depth 2202 may vary along the length 206 of the composite spar 202 or a portion of the length 206. Additionally, this same process may be performed on the sacrificial member 204 at flange 214, but the machining depth 2202 and / or initial thickness 310 may vary between flanges 214-215.

[0041] Another process that can be used to determine the amount of material to remove from the sacrificial member 204 after the composite assembly 200 has cured and hardened can be to statically define the machining depth 2202 based on the expected tolerances between the composite spar 202 and the skin panel 402 and / or skin panel 403, and then adjust the machining depth 2202 if gaps are found between the composite spar 202 and the skin panel 402 and / or skin panel 403.

[0042] First, the machining depth 2202 of the sacrificial member 204 can first be calculated (see step 1202 in FIG. 12 ). For example, components of the aircraft structure 400, during manufacturing, have manufacturing variations that can be estimated and used to generate worst-case scenarios for gaps between the flange 214 and the skin panel 402 and / or the flange 215 and the skin panel 403 of the aircraft structure 400 and can be used to initially determine the machining depth 2202. FIG. 22 shows an example of a calculated machining depth 2202 for the sacrificial member 204, illustrating how much material is removed from the outer surface 220 of the sacrificial member 204.

[0043] A machining process is performed on the outer surface 220 of the sacrificial member 204 to remove material to the machining depth 2202 shown in FIG. 22 (see step 1204), and the composite assembly 200 and skin panel 403 are reassembled as shown in FIG. 5 (see step 1206 of FIG. 12). After reassembly, the spacing between the contact surface 316 of the sacrificial member 204 and the inner surface 405 of the skin panel 403 is checked to determine whether a gap exists (see step 1208). If no gap exists, as shown in FIG. 5, the machining depth 2202 can be reused when machining subsequent composite assemblies (see step 1212). However, if a gap 2402 exists between the contact surface 316 and the inner surface 405 of the skin panel 403 at any point along the length 206 of the composite spar 202, as shown in FIG. 24, the machining depth 2202 may be recalculated (e.g., reduced) for machining subsequent composite spars to be manufactured to mitigate the gap 2402 (see step 1210).

[0044] Figure 25 is a flowchart of another method 2500 of manufacturing the composite assembly 200 in an exemplary embodiment, and Figure 26 is a flowchart illustrating additional details of the method 2500. Figures 27-28 are cross-sectional views of the flange 215 along section line CC in Figure 2 in an exemplary embodiment.

[0045] Step 2502 of method 2500 includes performing a first composite layup 1402 on a layup mandrel 1302 (see FIG. 14 ) that defines the contours of the composite spar 202. This step may be similar to step 602 of method 600 described above. Step 2504 includes performing a second composite layup 1502 on flange portion 1406 and / or flange portion 1407 of the first composite layup 1402, as shown in FIG. 15 . This step may be similar to step 604 of method 600 described above.

[0046] Step 2506 includes co-curing the first composite layup 1402 and the second composite layup 1502. This step may be similar to step 606 of the method 600 described above. The result of this process is the composite assembly 200 of FIG. 2.

[0047] Step 2508 includes calculating a machining depth of the sacrificial member 204 based on an estimated clearance tolerance between the flange 215 of the composite spar 202 and the skin panel 403 of the aircraft structure 400, and step 2510 includes machining the outer surface 220 of the sacrificial member 204 based on the machining depth.

[0048] In some cases, the machining depth 2702 is constant along the length 206 of the composite spar, as shown in FIG. 27. In other cases, the machining depth may vary along a portion of the length 206 of the composite spar 202, as shown in FIG. 28 (see step 2602 of FIG. 26). In FIG. 28, machining of the outer surface 220 of the sacrificial member 204 is performed to a first machining depth 2802 along a portion 2804 of the length 206 of the composite spar, and machining of the outer surface 220 of the sacrificial member 204 is performed to a second machining depth 2806 along a portion 2808 of the length 206 of the composite spar 202, but to different depths. As a result, the contact surface 316 of the sacrificial member 204 may vary in shape along the length 206 of the composite spar 202. Although the method 2500 is described with respect to the flange 215 and skin panel 403 of the composite spar 202 of the aircraft structure 400, the method 2500 applies equally to the flange 214 and skin panel 402 of the composite spar 202 of the aircraft structure 400, or any other type of structure for an aircraft that uses a spar.

[0049] The use of co-cured sacrificial member 204 on composite spar 202 integrates a sacrificial surface into flange 214 and / or flange 215, thereby eliminating the manual step of shimming the spar during the assembly process. In some cases, the machining of the sacrificial ply is performed based on estimated tolerances in the stackup of components that form the aircraft structure that uses the spar, thereby allowing the spar to be reused between different builds of the aircraft structure at the factory. In other cases, 3D scans may be performed on components of a specific assembly and linked to those specific components for assembly at the factory.

[0050] Embodiments of the disclosure may be described in the context of an aircraft manufacturing and service method 2900, as shown in FIG. 29 , and an aircraft 3004, as shown in FIG. 30 . Before production begins, the example method 2900 may include specification and design 2902 of the aircraft 3004, and material procurement 2904. During production, component and subassembly manufacturing 2906 and system integration 2908 of the aircraft 3004 occurs. The aircraft 3004 may then go through certification and delivery 2910 for placement in service 2912. While in service with a customer, the aircraft 3004 is scheduled for routine maintenance and service 2914 (which may also include modification, reconfiguration, refurbishment, etc.).

[0051] Each process of method 2900 may be performed or carried out 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 major system subcontractors; a third party may include, but is not limited to, any number of vendors, subcontractors, and suppliers; an operator may be an airline, a leasing company, a military entity, a service organization, etc.

[0052] 30 , an aircraft 3004 produced by exemplary method 2900 may include an airframe 3002 having a plurality of systems 3016 and an interior 3006. Examples of high-level systems 3016 include one or more of a propulsion system 3008, an electrical system 3010, a hydraulic system 3012, and an environmental system 3014. Any number of other systems may be included. While an aerospace example is shown, the principles described herein may be applied to other industries, such as the automotive industry.

[0053] Apparatus and methods embodied herein may be used during any one or more stages of manufacturing and service method 2900. For example, components or subassemblies corresponding to process 2906 may be fabricated or manufactured in a manner similar to components or subassemblies produced while aircraft 3004 is in service. Also, one or more apparatus embodiments, method embodiments, or combinations thereof may be used during component and subassembly manufacturing 2906 and system integration 2908, e.g., by substantially facilitating the assembly of, or substantially reducing the cost of, aircraft 3004. Similarly, one or more apparatus embodiments, method embodiments, or combinations thereof may be used during maintenance and service 2914 while aircraft 3004 is in service, for example, without limitation.

[0054] Although specific embodiments have been described herein, the scope is not limited to those specific embodiments, but rather is defined by the following claims and their equivalents. [Explanation of symbols]

[0055] 100 wings 102 Ribs 104 Front Spar 105 Rear Spar 106 leading edge 107 Trailing edge 108 Nose Rib 110 Outer Panel 200 Composite Assemblies 202 Composite Spar 204 Sacrificial Members 206 Length 208 Web 210 Main surface 212 Side of Web 208 213 Side of the Web 208 214, 215 flanges 216 Composite spar 202 width 218 Composite spar 202 height 220 outer surface of sacrificial member 204 302, 303 Bends 304 Outer surface of flange 214 305 Outer surface of flange 215 307 End of flange 214 307 End of flange 215 308 Edge 310 Initial thickness 314 Final Thickness 316 Contact surface 400 aircraft structure 402, 403 Skin Panel 404 Inner surface of skin panel 402 405 Inner surface of skin panel 403 406 area 502, 504 layers 600 ways 1302 Lay-up mandrel 1304 flat area 1306 First Side Area 1308 Second Side Area 1402 First Composite Layup 1404 Web section 1406, 1407 flange part 1408, 1409 Bends 1502 Second Composite Layup 1504 Edge of second composite layup 1502 1602 Adhesive film 1604 Work surface 1606 layers 1608 Thickness 1610 Exterior 2202 Machining depth 2402 gap 2500 ways 2702 Machining depth 2802 First machining depth 2804 parts 2806 Second Machining Depth 2808 parts 2900 method 2902 Specifications and Design 2904 Material Procurement 2906 Manufacture of components and subassemblies 2908 System Integration 2910 Authentication and Delivery 2912 in service 2914 Maintenance and inspection 3002 aircraft 3004 Aircraft 3006 Internal 3008 Propulsion System 3010 Electrical System 3012 Hydraulic System 3014 Environmental Systems 3016 High Level System

Claims

1. A composite spar, The web and a composite spar including: a flange projecting from a side of the web; a sacrificial member of composite material co-cured with said composite spar on an outer surface of at least one of said flanges; Equipped with the sacrificial member has an outer surface machined to match an inner surface of the at least one skin panel for an aircraft structure to form a contact surface with the at least one skin panel; a composite assembly, wherein the sacrificial member includes a tapered edge that reduces a thickness of the sacrificial member toward a bend in the composite spar between the web and at least one of the flanges.

2. the tapered edge forms a slope having a taper ratio of 1:1; The composite assembly of claim 1.

3. the sacrificial member and the composite spar comprise carbon fiber reinforced polymer (CFRP) plies; 3. A composite assembly according to claim 1 or 2.

4. an adhesive film disposed between at least one of the flanges and the sacrificial member; The composite assembly of any one of claims 1 to 3, further comprising:

5. the sacrificial member having a thickness of 0.076 cm to 0.203 cm; A composite assembly according to any one of claims 1 to 4.

6. 1. A method of manufacturing a composite assembly, comprising: assembling a first composite layup defining a web portion and a flange portion for a composite spar of the composite assembly; assembling a second composite layup over at least one of the flange portions to define a sacrificial member for the composite assembly; co-curing the first composite lay-up and the second composite lay-up to solidify the composite assembly; machining an outer surface of the sacrificial member to conform to an inner surface of at least one skin panel to form a contact surface for the at least one skin panel for an aircraft structure; Including, Assembling the second composite lay-up comprises: creating a tapered edge in the second composite lay-up; aligning the tapered edge with a bend in the first composite layup between the web portion and at least one of the flange portions such that a thickness of the second composite layup decreases toward the bend; The method further comprises:

7. generating the tapered edge comprises: Forming a lamp with a 1:1 taper ratio The method of claim 6, further comprising:

8. the first composite layup and the second composite layup include carbon fiber reinforced polymer (CFRP) plies; The method according to claim 6 or 7.

9. Assembling the second composite lay-up comprises: assembling the second composite lay-up onto an adhesive film; transferring the second composite layup to at least one of the flange portions with the adhesive film in contact with at least one of the flange portions; The method of any one of claims 6 to 8, further comprising:

10. Assembling the second composite lay-up comprises: assembling the second composite lay-up onto an adhesive film; cutting a tapered edge of the second composite lay-up; transferring the second composite layup to at least one of the flange portions with the adhesive film in contact with at least one of the flange portions; aligning the tapered edge with a bend in the first composite layup between the web portion and at least one of the flange portions such that a thickness of the second composite layup decreases toward the bend; The method of any one of claims 6 to 9, further comprising:

11. machining the outer surface of the sacrificial member; identifying the at least one skin panel of an aircraft structure to be attached to the composite spar; calculating interference between the inner surface of the at least one skin panel and the outer surface of the sacrificial member; machining the outer surface of the sacrificial member based on the interference to mate the outer surface of the sacrificial member with the inner surface of the at least one skin panel to form the contact surface of the at least one skin panel; The method of any one of claims 6 to 10, further comprising:

12. machining the outer surface of the sacrificial member; calculating a machining depth of the sacrificial member along a length of the composite spar based on an estimated spacing tolerance between the composite spar and the at least one skin panel; machining the outer surface of the sacrificial member based on the calculated machining depth; assembling the composite spar and the at least one skin panel to form the aircraft structure; determining that a gap exists between the contact surface of the sacrificial member and the inner surface of the at least one skin panel; modifying the machining depth to mitigate the gap in subsequent machining processes for the sacrificial member for the composite spar; The method of any one of claims 6 to 11, further comprising:

13. 1. A method of manufacturing a composite assembly, comprising: performing a first composite layup on a layup mandrel that defines a composite spar profile of the composite assembly; performing a second composite layup on a flange portion of the first composite layup defined by a contour; co-curing the first composite layup and the second composite layup to solidify the composite assembly, the first composite layup forming the composite spar and the second composite layup forming a sacrificial member on a flange on the composite spar; calculating a machining depth of the sacrificial member based on an estimated clearance tolerance between the flange of the composite spar and at least one skin panel for an aircraft structure; machining an outer surface of the sacrificial member along at least a portion of a length of the composite spar based on the machining depth to form a contact surface for the at least one skin panel; Including, performing the second composite layup; conducting the second composite lay-up on an adhesive film; cutting a tapered edge of the second composite lay-up; transferring the second composite lay-up to the flange portion of the first composite lay-up with the adhesive film in contact with the flange portion; aligning the tapered edge of the second composite layup with a bend in the first composite layup between a web portion and a flange portion defined by the contour such that a thickness of the second composite layup decreases toward the bend; A method comprising:

14. cutting the tapered edge Forming a lamp with a 1:1 taper ratio 14. The method of claim 13, further comprising:

15. the first composite layup and the second composite layup include carbon fiber reinforced polymer (CFRP) plies; 15. The method of any one of claims 13 or 14.

16. machining the outer surface Varying the machining depth of the sacrificial member along at least a portion of the length of the composite spar. The method according to any one of claims 13 to 15, comprising:

17. determining that a gap exists between the contact surface of the sacrificial member and an inner surface of the at least one skin panel; modifying the machining depth to mitigate the gap in subsequent machining processes for the sacrificial member for the composite spar; The method of any one of claims 13 to 16, further comprising:

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