Composite spars with integrated sacrificial surfaces

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 that utilize spars as structural members.

Background Art

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

[0003] To ensure proper fitting of the skin panel to the spar, shims can be installed around the spar where the inner surface of the skin panel contacts the spar. The shims are used to eliminate possible machining processes on the spar itself that could have an adverse effect on the structural integrity of the spar.

[0004] Generally, installing shims on the spar is a time-consuming manual operation, and additional labor is required for the assembly time and disassembly time of the wing and other aircraft structures using the spar to insert the shims into the gap between the skin panel and the spar.

Summary of the Invention

Problems to be Solved by the Invention

[0005] Therefore, based on the foregoing description, it is still desirable to improve the aircraft manufacturing process, particularly the process for assembling aircraft structures including spars.

Means for Solving the Problems

[0006] A composite assembly is described that includes a composite spar, in which one or more sacrificial members on a flange are co-cured to form a unified sacrificial surface for the composite spar. Generally, a spar is an elongated structure that includes a web defining the main surface of the spar and flanges projecting from the sides of the web. The spar extends in the wingspan direction within the wing perpendicular or nearly perpendicular to the fuselage and forms a major structural component of the wing. The web within the spar is positioned perpendicularly within the wing, and the flange defines a surface for attaching the skin panel. During assembly, a gap may exist between the flange and the skin panel. In the embodiments described herein, the sacrificial members co-cured with the flange are machined to match the outer surface of the sacrificial members to the surface of the skin panel, providing a technical benefit of reducing the gap between the spar and the skin panel in a wing or other aircraft structure utilizing a spar.

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

[0008] Another embodiment includes a method for manufacturing a composite assembly. This method includes the steps of: assembling a first composite layup defining web and flange portions for a composite spar of the composite assembly; assembling a second composite layup on 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. This method further includes the step of machining the outer surface of the sacrificial member to match the inner surface of at least one skin panel for an aircraft structure in order to form a contact surface for at least one skin panel.

[0009] Another embodiment includes a method for manufacturing a composite assembly. This method includes the steps of: performing a first composite layup on a layup mandrel defining the outer shape of a composite spar of the composite assembly; performing a second composite layup on the flange portion of the first composite layup defined by the outer shape; and co-curing the first and second composite layups to solidify the composite assembly, wherein the first composite layup forms a composite spar and the second composite layup forms a sacrificial member on the flange of the composite spar. This method further includes the steps of: calculating a machining depth of the sacrificial member based on an estimated spacing tolerance between the flange of the composite spar and at least one skin panel for an aircraft structure; and machining the 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 for at least one skin panel.

[0010] The features, functions, and advantages described above can be achieved independently in various embodiments or combined in even more embodiments, and further details can be understood by referring to the following description and drawings.

[0011] Herein, several embodiments are described, merely as examples, with reference to the attached drawings. The same reference numerals represent the same element or element of the same type in all drawings. [Brief explanation of the drawing]

[0012] [Figure 1] This shows the wing of an aircraft using conventional technology. [Figure 2] This is an isometric view of a composite assembly in an exemplary embodiment. [Figure 3] This is a cross-sectional view of the composite assembly along the cutting line AA in Figure 2 in an exemplary embodiment. [Figure 4] A portion of the aircraft structure in an exemplary embodiment is shown. [Figure 5] This is a diagram of the region shown in Figure 4 in an exemplary embodiment. [Figure 6] A method for manufacturing a composite structure in an exemplary embodiment. [Figure 7] Shows further details of the method of FIG. 6 in an exemplary embodiment. [Figure 8] Shows further details of the method of FIG. 6 in an exemplary embodiment. [Figure 9] Shows further details of the method of FIG. 6 in an exemplary embodiment. [Figure 10] Shows further details of the method of FIG. 6 in an exemplary embodiment. [Figure 11] Shows further details of the method of FIG. 6 in an exemplary embodiment. [Figure 12] Shows further details of the method of FIG. 6 in an exemplary embodiment. [Figure 13] An isometric view of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 14] An isometric view of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 15] An isometric view of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 16] An isometric view of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 17] An isometric view of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 18] An isometric view of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 19] An isometric view of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 20] An isometric view of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 21] An isometric view of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 22] An isometric view of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 23]An isometric view of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 24] An isometric view of various stages of the manufacturing method of FIG. 6 in an exemplary embodiment. [Figure 25] A flowchart of another method for manufacturing a composite structure in an exemplary embodiment. [Figure 26] Shows further details of the method of FIG. 25 in an exemplary embodiment. [Figure 27] A cross-sectional view of the composite assembly along the cutting line C-C of FIG. 2 in an exemplary embodiment. [Figure 28] A cross-sectional view of the composite assembly along the cutting line C-C of FIG. 2 in an exemplary embodiment. [Figure 29] A flowchart showing a method for manufacturing and maintaining an aircraft in an exemplary embodiment. [Figure 30] A schematic view of an aircraft in an exemplary embodiment.

Mode for Carrying Out the Invention

[0013] The drawings and the following description show specific exemplary embodiments. It will be understood by those skilled in the art that, although not explicitly described or illustrated herein, various configurations that embody the principles described herein and fall within the intended scope of the appended claims can be devised. Further, any examples described herein are intended to assist in understanding the principles of the present disclosure and should not be construed as limiting. 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.

[0014] Figure 1 shows a prior art wing 100. The wing 100 includes a number of ribs 102 extending between a front spar 104 and a rear spar 105, with the front spar 104 located near the leading edge 106 of the wing 100 and the rear spar 105 extending toward the trailing edge 107 of the wing 100. In Figure 1, the ribs 102 are spaced apart along the length of the wing 100, and their periphery forms the airfoil shape of the wing 100. In Figure 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 leading edge shape of the airfoil at the leading edge 106. The wing 100 also includes a skin 110, partially removed in this figure, to show the structure beneath the wing 100.

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

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

[0017] Composite parts such as carbon fiber reinforced polymer (CFRP) parts and / or glass fiber reinforced polymer (GFRP) parts are initially constructed by stacking multiple material layers, sometimes called reinforcement layers. While the individual fibers within each layer are aligned parallel to one another, different layers exhibit different fiber orientations along different dimensions to enhance the strength of the resulting composite part. These layers are pre-impregnated with a matrix material, such as an uncured thermosetting resin or thermoplastic resin, and are called "prepregs." Alternatively, the layers may be laminated in a dry state (i.e., "dry fibers"), and then the matrix material may be injected before curing.

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

[0019] In this embodiment, the sacrificial member 204 is co-cured with the flanges 214 and / or 215. After the composite assembly 200 has cured, the outer surface 220 of the sacrificial member 204 can be machined as needed during the manufacturing process to reduce any gaps that may occur between the flanges 214 and / or 215 of the composite spar 202 and other aircraft structures such as skin panels. The use of the sacrificial member 204 reduces the use of manual shims that would normally be introduced when spar flanges are fitted into skin panels and gaps exist, thereby providing the technical benefit of reducing the time and effort required to assemble various aircraft structures that use spars and skin panels.

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

[0021] In this embodiment, the sacrificial member 204 is shown on both flanges 214-215, but in some embodiments, the 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 along the cutting line AA of Figure 2 in an exemplary embodiment. In this embodiment, the composite spar 202 includes a bend 302 between the web 208 and the flange 214 at the side 212 of the web 208, and the sacrificial member 204 is positioned 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 adjacent to the bend 302. 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 a ramp. Using 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 later.

[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. Generally, the initial thickness 310 is pre-selected based on assumed tolerances between the flange 214 and other components on the aircraft, such as skin panels. After co-curing the composite assembly 200, the outer surface 220 of the sacrificial member 204 may be machined (for example, material is removed from the outer surface 220 of the sacrificial member 204 along the length 206 or portion of the composite spar 202 (see Figure 2)), thereby reducing the initial thickness 310 of the sacrificial member 204 to a final thickness 314 based on the machining profile or machining depth. Machining of the sacrificial member 204 also creates a contact surface 316 with the skin panel. Although not shown in Figure 3, both the composite spar 202 and the sacrificial member 204 are formed from a composite material such as CFRP ply.

[0024] Similar to the flange 214, the composite spar 202 includes a bend 303 between the web 208 and the flange 215 on the side surface 213 of the web 208, and the sacrificial member 204 is positioned on the outer surface 305 of the flange 215. In this embodiment, the sacrificial member 204 extends from the end 307 of the flange 215 toward the bend 303 and terminates at an edge 308 adjacent to 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 tapered or ramped for similar reasons as 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. Generally, the initial thickness 310 is selected based on assumed tolerances between the flange 215 and other components on the aircraft, such as skin panels, and then the outer surface 220 of the sacrificial member 204 can be machined (for example, material is removed from the outer surface 220 of the sacrificial member 204 at the flange 215 along the length 206 of the composite spar 202 (see Figure 2)), thereby reducing the initial thickness 310 of the sacrificial member 204 to the final thickness 314 based on the machined profile. Furthermore, the amount of material removed from the sacrificial member 204 may vary along the length 206 or a portion of the length 206 of the composite spar 202, and / or depending on whether the sacrificial member 204 is located in the flange 214-215.

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

[0026] As mentioned above, shims were often manually placed between the spar flange and the skin panel during the assembly process to compensate for or reduce the gap formed between these two components. In general, direct machining on the composite spar 202 is undesirable because the machining process can remove the fiber layer and impair the structural integrity of the composite spar 202. The use of shims makes direct machining on the spar impossible, but 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 components, and joining the shims to the spar flange. The shims can then be machined to achieve the final fit between the spar flange and the skin panel. The final assembly of the spar and skin panel can then be carried out.

[0027] With the sacrificial member 204 co-hardened with the flanges 214 and / or 215 of the composite spar 202, a machining process can be performed on the outer surface 220 of the sacrificial member 204 before assembling the aircraft structure 400, thereby saving time and effort compared to the conventional manual shim process. For example, the machined profile or machining depth of the sacrificial member 204 can be generated based on several different factors, including assumed tolerances between the flanges 214-215 of the composite spar 202 and the inner surfaces 404-405 of the skin panels 402-403. Once the machined profile or machining depth is selected and the composite assembly 200 is tested for fitting to 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 reduce the gaps in future builds of the aircraft structure 400.

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

[0029] Figure 5 shows a diagram of region 406 in Figure 4 in an exemplary embodiment. In this embodiment, the edge 308 of the sacrificial member 204 is shaped like a ramp or taper adjacent to the bend 303 of the composite spar 202. The use of a taper or ramp on the edge 308 may be used to mitigate deformation of the composite spar 202 that may occur during the manufacturing process, which may cause waves in the fiber layers constituting the composite spar 202 due to vacuum compression during curing. The edge 308 may be cut into a ramp shape after layup, or it may be laminated as a ramp during manufacturing. When the edge 308 is laminated as a ramp, the width of the layer may decrease as the distance from the outer surface 305 of the flange 215 increases (for example, by making the layer 502 adjacent to the outer surface 305 of the flange 215 wider than the layer 504 distal to the outer surface 305 of the flange 215). Figure 5 also shows the contact surface 316 that contacts the inner surface 405 of the skin panel 403, which is generated after machining the outer surface 220 (see Figure 3) of the sacrificial member 204 until the final thickness 314 is achieved.

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

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

[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. Figure 14 shows the first composite layup 1402 on the layup mandrel 1302, defining the web portion 1404, flange portions 1406-1407, and bend portions 1408-1409 of the composite spar 202 defined by the shape or outline beneath the layup mandrel 1302. Generally, the first composite layup 1402 includes continuous fiber layers formed as a prepreg or dry layup, where the orientation may vary from layer to layer. Furthermore, the layers may be continuous from the end 306 of flange 214 to the end 307 of flange 215. As previously mentioned, the first composite layup 1402 may include a BMS8-276 tape having a pseudo-isotropic layup with very high rigidity after curing.

[0033] Step 604 includes assembling the second composite layup 1502 (see Figure 15) on the flange portions 1406 and / or 1407 of the first composite layup 1402. For example, a flat layup may be performed on the second composite layup 1502 and transferred to or in direct contact with the flange portions 1406 and / or 1407 of the first composite layup 1402. In another example, the AFP machine can perform the layup directly on the flange portions 1406 and / or 1407 of the first composite layup 1402. As previously stated, the second composite layup 1502 may include a GFRP material with moderate post-curing stiffness compared to BMS8-276 tape and BMS8-276 fabric laminated with a changing + / - 45 degree orientation, or a GFRP material with lower post-curing stiffness compared to 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 solidify the composite assembly 200. For example, the first composite layup 1402 and the second composite layup 1502 may be placed in a bag, pressure may be applied to the first composite layup 1402 and the second composite layup 1502 under vacuum, and heating may be performed to solidify the composite assembly 200. The structure obtained after curing is a composite assembly 200 as shown in Figure 2, in which the composite spar 202 has a web 208 formed from web portion 1404 and flanges 214-215 formed from flange portions 1406-1407 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 match the surface of one or more skin panels. For example, machining of the outer surface 220 of the sacrificial member 204 is performed to remove material from the sacrificial member 204 and to form the contact surface 316 of the skin panel 402 and / or skin panel 403. Prior to the machining process, the initial thickness 310 of the sacrificial member 204 may be about 0.08 inches to 0.12 inches. After the machining process, the final thickness 314 of the sacrificial member 204 may be about 0.03 inches to 0.08 inches.

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

[0037] If the 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 the adhesive film 1602 placed on the work surface 1604, as shown in Figure 16 (see step 902 in Figure 9). As shown in Figure 16, the second composite layup 1502 includes a number of vertically laminated layers 1606 to achieve the desired thickness 1608 of the second composite layup 1502, which is used to define the initial thickness 310 after curing (see Figure 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 has cured.

[0038] After assembling the second composite layup 1502 onto the adhesive film 1602, the second composite layup 1502 and the adhesive film 1602 are transferred to the flange portions 1406 and / or 1407 of the first composite layup 1402 during the assembly process, as shown in Figure 17 (see step 904), and the adhesive film 1602 comes into contact with the outer surfaces 304 of flange 214 and / or 305 of flange 215. To produce the composite assembly 200 of Figure 18, vacuum bagging, compression, and curing can be performed with the adhesive film 1602 positioned between the sacrificial member 204 and flange 214 and / or flange 215. Figure 19 shows a cross-sectional view of Figure 18 along the cutting line BB, showing the position of the adhesive film 1602 between the sacrificial member 204 and flange 214 and / or flange 215.

[0039] If the edge 308 is tapered to the sacrificial member 204, the second composite layup 1502 can be assembled onto the adhesive film 1602 as described above with respect to Figures 9 and 16, and the tapered edge 1504 of the second composite layup 1502 can be cut (see step 1002 and Figure 20), and then, with the edge 1504 aligned to the bend 1408 and / or bend 1409 (see step 1004 and Figure 17), the second composite layup 1502 can be transferred to the flange portion 1406 and / or flange portion 1407 of the first composite layup 1402 (see step 904). 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 Figure 21.

[0040] As described above, there are several processes that can be used to determine how much material to remove from the outer surface 220 of the sacrificial member 204 after the composite assembly 200 has hardened and solidified. One process involves connecting or associating the composite assembly 200 with specific skin panels 402-403 of the aircraft structure 400 (see Figure 4). First, one or more skin panels to be connected or mated to the composite spar 202 can be identified (see step 1102 in Figure 11). Depending on the design of the structure using the composite spar 202, multiple skin panels may be bonded or mated to the composite spar 202 along their length 206. For example, skin panel 403 may be identified as being assigned to be installed on the flange 215 of the composite spar 202 (see Figure 5). The interlock between skin panel 403 and the outer surface 220 of the sacrificial member 204 can be calculated (see step 1104). For example, 3D scans of skin panel 403 and the composite assembly 200 may be performed and used to calculate the interlock. Next, the outer surface 220 of the sacrificial member 204 can be machined based on the interlocking allowance (see step 1106). For example, the outer surface 220 of the sacrificial member 204 may have an initial thickness 310 before machining, as shown in Figure 22, which is based on the thickness 1608 of the second composite layup 1502 before curing (see Figure 16). Then, before assembling the composite spar 202 and the skin panel 403 together, the outer surface 220 can be machined to a final thickness 314, as shown in Figure 22. The machining depth 2202 cutting 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. Although Figure 22 simply shows a cross-section of a particular part of the aircraft structure 400 for illustrative purposes, the machining depth 2202 may vary along the length 206 or a portion of the length 206 of the composite spar 202. Furthermore, this same process may be carried out on the sacrificial member 204 at the flange 214, but the machining depth 2202 and / or initial thickness 310 may differ between flanges 214 and 215.

[0041] Another process that can be used to determine the amount of material to be removed from the sacrificial member 204 after the composite assembly 200 has hardened and solidified is to statically define the machining depth 2202 based on the assumed tolerance between the composite spar 202 and the skin panel 402 and / or skin panel 403, and then adjust the machining depth 2202 if a gap is 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 be calculated (see step 1202 in Figure 12). For example, the components of the aircraft structure 400 have manufacturing variations that can be estimated and used during manufacturing to generate a worst-case scenario for the gap between the flange 214 and skin panel 402 and / or flange 215 and skin panel 403 of the aircraft structure 400, and which can be used to initially determine the machining depth 2202. Figure 22 shows an example of a calculated machining depth 2202 for the sacrificial member 204, showing how much material is removed from the outer surface 220 of the sacrificial member 204.

[0043] The machining process is performed on the outer surface 220 of the sacrificial member 204 to remove material to a machining depth 2202 shown in Figure 22 (see step 1204), and the composite assembly 200 and skin panel 403 are reassembled as shown in Figure 5 (see step 1206 in Figure 12). After reassembly, the gap between the contact surface 316 of the sacrificial member 204 and the inner surface 405 of the skin panel 403 is checked to determine if a gap exists (see step 1208). If no gap exists, as shown in Figure 5, the machining depth 2202 can be reused when machining the subsequent composite assembly (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 Figure 24, the machining depth 2202 may be recalculated (e.g., reduced) to machine the subsequent composite spar to be manufactured in order to mitigate the gap 2402 (see step 1210).

[0044] Figure 25 is a flowchart of another method 2500 for manufacturing the composite assembly 200 in an exemplary embodiment, and Figure 26 is a flowchart showing additional details of method 2500. Figures 27-28 are cross-sectional views of the flange 215 along the cutting 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 Figure 14) that defines the outline of the composite spar 202. This step may be the same as step 602 of Method 600 described above. Step 2504 includes performing a second composite layup 1502 on flange portions 1406 and / or 1407 of the first composite layup 1402, as shown in Figure 15. This step may be the same as 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 the same as step 606 of the previously described method 600. The result of this process is the composite assembly 200 shown in Figure 2.

[0047] Step 2508 includes calculating the machining depth of the sacrificial member 204 based on the estimated spacing 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 Figure 27. In other cases, the machining depth may vary along a portion of the length 206 of the composite spar 202, as shown in Figure 28 (see step 2602 in Figure 26). In Figure 28, machining of the outer surface 220 of the sacrificial member 204 is carried out along a portion 2804 of the length 206 of the composite spar to a first machining depth 2802, and machining of the outer surface 220 of the sacrificial member 204 is carried out along a portion 2808 of the length 206 of the composite spar 202 to a second machining depth 2806, but to different depths. As a result, the contact surface 316 of the sacrificial member 204 may change shape along the length 206 of the composite spar 202. Although Method 2500 is described in relation to the flange 215 and skin panel 403 of the composite spar 202 of the aircraft structure 400, Method 2500 is equally applicable to the flange 214 and skin panel 402 of the composite spar 202 of the aircraft structure 400, or to any other type of structure for aircraft that uses a spar.

[0049] The use of co-hardened sacrificial members 204 on the composite spar 202 integrates the sacrificial surface with flanges 214 and / or 215, thereby eliminating the manual step of attaching shims to the spar during the assembly process. In some cases, the machining of the sacrificial ply is carried out based on estimated tolerances in the lamination of the components forming the aircraft structure using 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 the components of a particular assembly and linked to these specific components for assembly at the factory.

[0050] Embodiments of this disclosure can be described in relation to an aircraft manufacturing and maintenance method 2900 as shown in Figure 29 and an aircraft 3004 as shown in Figure 30. Before the commencement of manufacturing, exemplary method 2900 may include the specification and design 2902 of the aircraft 3004, and material procurement 2904. During manufacturing, the manufacturing of components and subassemblies 2906 of the aircraft 3004, and system integration 2908 are carried out. The aircraft 3004 can then undergo certification and transport 2910 to be deployed in service 2912. During the aircraft's service by the customer, a schedule is set for periodic maintenance and inspection 2914 of the aircraft 3004 (which may also include modifications, reconfigurations, refurbishments, etc.).

[0051] Each process of Method 2900 may be carried out or performed by a system integrator, a third party, and / or an operator (e.g., a customer). For the purposes of this description, the system integrator may include, but is not limited to, any number of aircraft manufacturers and major system subcontractors; the third party may include, but is not limited to, any number of vendors, subcontractors, and suppliers; and the operator may be an airline, leasing company, military entity, service organization, etc.

[0052] As shown in Figure 30, an aircraft 3004 manufactured by exemplary method 2900 may include a fuselage 3002 having multiple systems 3016 and internal components 3006. An example of a high-level system 3016 includes one or more of the propulsion system 3008, electrical system 3010, hydraulic system 3012, and environmental system 3014. Any number of other systems may be included. Although an aerospace example is shown, the principles described herein can be applied to other industries such as the automotive industry.

[0053] The apparatus and methods embodied herein may be used in any one or more stages of the manufacturing and maintenance methods 2900. For example, components or subassemblies corresponding to process 2906 may be manufactured or produced in the same manner as components or subassemblies produced during the operation of the aircraft 3004. Also, one or more embodiments of the apparatus, embodiments of the methods, or combinations thereof may be used in the manufacturing of components and subassemblies 2906 and system integration 2908, for example, by substantially accelerating the assembly of the aircraft 3004 or substantially reducing the cost of the aircraft. Similarly, one or more embodiments of the apparatus, embodiments of the methods, or combinations thereof may be used in maintenance and inspection 2914, for example, but not limited to, while the aircraft 3004 is in operation.

[0054] While specific embodiments have been described herein, the scope is not limited to those specific embodiments. Rather, the scope is defined by the following claims and equivalents. [Explanation of Symbols]

[0055] 100 wings 102 Rib 104 Front Spar 105 Rear sparring 106 Leading edge 107 Trailing edge 108 Nose Rib 110 Outer panel 200 composite assemblies 202 Combined Spar 204 Sacrificial member 206 Length 208 Web 210 Main surface 212 Side of Web 208 213 Side of Web 208 214, 215 flange 216 Composite spar width 202 218 Height of composite spar 202 220 Outer surface of sacrificial member 204 302, 303 Bending section 304 Flange 214 outer surface 305 Flange 215 outer surface 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 Skin panel 403 interior 406 area 502, 504 layers 600 ways 1302 Layup Mandrel 1304 flat area 1306 First lateral region 1308 Second lateral region 1402 First composite layup 1404 Web section 1406, 1407 Flange section 1408, 1409 Bending section 1502 Second composite layup 1504 Edge of the 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 Manufacturing of components and subassemblies 2908 System Integration 2910 Authentication and Transport 2912 in service 2914 Maintenance and inspection 3002 aircraft 3004 Aircraft 3006 Internal 3008 Propulsion System 3010 Electrical Systems 3012 Hydraulic System 3014 Environmental Systems 3016 High-Level Systems

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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