Multi-segment spar manufacturing

Abstract

The invention relates to multi-section spar manufacturing. Systems and methods for manufacturing multi-section spars for aircraft are provided. In one example, the method includes: manufacturing preforms of fiber-reinforced material for spar sections; hardening the preforms to form spar sections; and joining the spar sections together to form a complete spar part. In addition to joining, other examples include co-curing spar sections and fastening spar sections. In further examples, the spar sections include described kinks or sub-kinks.

Description

TECHNICAL FIELD

[0001] The present disclosure relates to the field of manufacturing, in particular to the manufacturing of composite parts for aircraft. BACKGROUND

[0002] A spar of an aircraft extends within a wing from an inboard direction to an outboard direction and provides structural strength to the wing. The spar of the aircraft can be made, for example, of composite material. Because the wing is a particularly long component of the aircraft, multiple spars extend the length of the wing. The large size of the spars increases the complexity and expense of the equipment dedicated to laying up, curing, hardening, assembling the spars, and assembling the wing.

[0003] EP2186622 A1, according to its abstract, sets forth a composite structural member comprising a first composite section and a second composite section spliced together by an overlapping composite splice member. US10836121B2, according to its abstract, describes, in one aspect, a method of manufacturing a composite skin for a tiltrotor aircraft, the method comprising: providing a first skin in a mold, the first skin having a perimeter defined by a front edge, a rear edge, and an outboard end; providing a plurality of honeycomb panels on the first skin, the plurality of honeycomb panels having a large cell array, each cell having a width of at least 1 cm; assembling the plurality of honeycomb panels along a longitudinal axis of the first skin to form a honeycomb core having an outer perimeter that lies within the perimeter of the first skin; positioning a second skin on the honeycomb core, the second skin having an outer perimeter that lies within the perimeter of the first skin; and curing an adhesive to form a bond between the first skin, the honeycomb core, and the second skin to form the composite skin.

[0004] WO2010 / 122325 A1, according to its abstract, sets forth a structure comprising a cured composite part formed from a series of layers of fiber-reinforced composite material, a reinforcement plate attached to the composite part by an array of sharp prongs that partially penetrate the composite part, and a hole through the reinforcement plate and the composite part. The interface plate carries the array of prongs on a first side and is attached to the reinforcement plate on a second side.

[0005] EP3650333 A1, according to its abstract, sets forth an example aircraft wing comprising a first skin, a second skin opposite the first skin, and a composite spar. The composite spar comprises a double-flanged spar cap, a single-flanged spar cap, a spar web connecting the double-flanged spar cap and the single-flanged spar cap, and a tear strip. The double-flanged spar cap comprises an inward-facing flange and a first outward-facing flange, and the inward-facing flange and the first outward-facing flange are integral with the first skin during a co-curing process. The single-flanged spar cap comprises a second outward-facing flange attached to the second skin. The tear strip is stitched to an inboard side of the spar web along at least a portion of a length of the composite spar.

[0006] US 2016 / 121589 Al, according to its abstract, sets forth providing a method and related system for forming a composite material, including providing a composite fill having a release film applied to a top surface and a bottom surface and positioned on a contoured forming tool such that a portion of the composite fill overhangs a second upper surface of the forming tool. A shimming bar is abutted against the forming tool, the shimming bar having an uppermost surface that is positioned above the second upper surface of the forming tool. The combination of the positioning of the shimming bar and the use of the release film in contact with the composite fill allows the composite fill to follow the contoured forming tool without bridging or wrinkling the composite fill when using a forming film and applying a vacuum.

[0007] The problems discussed above are further complicated by the fact that many spars exhibit complex curvatures that correspond to the complex curvatures of the wing, which further complicates the manufacturing process.

[0008] Accordingly, it would be desirable to have methods and apparatus that take into account at least some of the problems discussed above as well as possibly other issues. SUMMARY

[0009] Embodiments described herein provide spars formed from a plurality of segments that are structurally integrated together via co-curing, co-bonding, or fastener installation. By assembling a spar with a plurality of segments, the size of the equipment used to manufacture the spar can be reduced. Moreover, because the segments are smaller than the entire spar, each segment can exhibit simpler curvatures (e.g., more linear profiles) than the resulting spar, which further reduces the complexity of the equipment used for layup, curing, etc.

[0010] In one aspect, a method for manufacturing a spar detail for an aircraft is provided. The method includes manufacturing a preform of fiber reinforced material for a spar segment, hardening the preform to form the spar segment, and bonding the spar segments together to form a complete spar detail.

[0011] In another aspect, a spar for an aircraft is provided. The spar includes a first spar segment including fiber reinforced material, the first spar segment including a splice region, a second spar segment including fiber reinforced material, the second spar segment including a first splice region and disposed in series with the first spar segment, and a splice reinforcement covering at least a portion of the first splice region of the second spar segment and the splice region of the first spar segment, the splice reinforcement bonded to the first spar segment and the second spar segment.

[0012] In yet another aspect, a method for manufacturing a spar part for an aircraft is provided. The method includes: manufacturing a preform of fiber-reinforced material for a first spar segment and a second spar segment; splicing an end of the preform of the first spar segment to an end of the preform of the second spar segment to define a splice region; applying at least one preform for a splice reinforcement to the splice region; and simultaneously hardening the preforms for the spar segments and the splice reinforcement to form a portion of the spar part.

[0013] In yet another aspect, a spar part is provided. The spar part includes: a preform for a first spar segment, the first spar segment including a splice region; a preform for a second spar segment, the second spar segment including a splice region, the splice regions arranged in series with one another in a spliced relationship; and a splice reinforcement preform, the preforms and the splice reinforcement preform being simultaneously cured, the splice reinforcement preform covering at least a portion of the splice regions to form a portion of the spar part.

[0014] In yet another aspect, a method for manufacturing a spar part for an aircraft is provided. The method includes: manufacturing a preform of fiber-reinforced material for a spar segment; hardening the preform to form the spar segment; and applying fasteners that couple the spar segments together to form a complete spar part.

[0015] In yet another aspect, a spar part for an aircraft is provided. The spar part includes: a first spar segment, the first spar segment including a fiber-reinforced material; a second spar segment, the second spar segment including a fiber-reinforced material and arranged in series with the first spar segment; a splice reinforcement, the splice reinforcement covering a splice region between the first spar segment and the second spar segment; and fasteners, the fasteners installed through the splice reinforcement, the first spar segment, and the second spar segment to form at least a portion of the spar part.

[0016] In another aspect, a method for manufacturing a spar for an aircraft is provided. The method includes: manufacturing preforms of fiber-reinforced material for spar segments, at least one of the spar segments including a kink, each kink being entirely contained within a preform; hardening the preforms to form the spar segments; and assembling the spar segments together to form a complete spar part exhibiting at least one of the kinks.

[0017] In yet another aspect, an aircraft spar is provided. The aircraft spar includes a first spar segment including a fiber-reinforced material and at least one splice region, a second spar segment including a fiber-reinforced material, at least one splice region, and a jog outside the splice region, where the respective splice regions are disposed in series with one another, and a splice reinforcement covering at least a portion of the splice region of the first spar segment and at least a portion of the corresponding splice region of the second spar segment, the splice reinforcement coupled to the first spar segment and the second spar segment.

[0018] In yet another aspect, a method for manufacturing a spar part of an aircraft is provided. The method includes manufacturing a preform for a first spar segment, the preform including a sub-jog proximate one end of the first spar segment, manufacturing a preform for a second spar segment, the preform including a sub-jog proximate one end of the second spar segment, aligning the ends of the preforms such that the sub-jogs are proximate to one another within a splice region, and joining the spar segments together in the splice region to form at least a portion of a spar part exhibiting a jog.

[0019] In another aspect, a spar part for an aircraft is provided. The spar part includes a first spar segment including a fiber-reinforced material and including a sub-jog disposed at an end, a second spar segment including a fiber-reinforced material and including a sub-jog disposed at an end, the end of the first spar segment having the sub-jog adjacent to the end of the second spar segment having the sub-jog such that the sub-jogs together form a jog, and the ends defining a splice region, and a splice reinforcement structurally integrating the first spar segment and the second spar segment within the splice region.

[0020] Other illustrative embodiments (e.g., methods and computer readable media related to the aforementioned embodiments) can be described below. The features, functions, and advantages described can be implemented independently in various embodiments of the present disclosure or combinations thereof and further details of these and other embodiments can be seen in the description below and in the drawings. BRIEF DESCRIPTION OF DRAWINGS

[0021] Some embodiments of the present disclosure will now be described, by way of example only, with reference to the attached drawings. In all the drawings, like reference numerals refer to like elements or elements of a like type.

[0022] Figure 1 is a perspective view of an aircraft including a fully assembled wing in an illustrative embodiment.

[0023] Figure 1A is a block diagram of a spar part in an illustrative embodiment.

[0024] Figure 2 A system for forming a spar part from multiple segments in an illustrative embodiment is shown.

[0025] Figure 3 is a drawing depicting a spar part formed from multiple spar segments in an illustrative embodiment.

[0026] Figure 4 is a flowchart showing a method for manufacturing a spar part from spar segments via co-bonding in an illustrative embodiment.

[0027] Figure 5 is a flowchart showing another method for manufacturing a spar part from spar segments via co-curing in an illustrative embodiment.

[0028] Figure 6 is a flowchart showing another method for manufacturing a spar part from spar segments via fasteners in an illustrative embodiment.

[0029] Figure 7 is a flowchart showing a method for manufacturing a spar part with a kink in an illustrative embodiment.

[0030] Figure 8 is a flowchart showing another method for manufacturing a spar part with a kink in an illustrative embodiment.

[0031] Figure 9 An aircraft wing including a spar in an illustrative embodiment is depicted.

[0032] Figure 9A A cross-section of a wing in an illustrative embodiment is depicted.

[0033] Figures 10-11 A spar part in an illustrative embodiment is depicted.

[0034] Figure 12 A kink at a portion of a beam detail in an illustrative embodiment is depicted.

[0035] Figures 13-14 An assembly of a kink that has been subdivided into bends in an illustrative embodiment is depicted.

[0036] Figures 15-17 A flat charge is formed into a preform having a predetermined cross-section in an illustrative embodiment is depicted.

[0037] Figure 18 A scarf joint between spar segments in an illustrative embodiment is depicted.

[0038] Figure 19 is a flowchart of a method of aircraft production and maintenance in illustrative embodiments.

[0039] Figure 20 is a block diagram of an aircraft in illustrative embodiments. DETAILED DESCRIPTION

[0040] The accompanying drawings and following description provide specific illustrative embodiments of the present disclosure. Thus, it will be apparent to one of ordinary skill in the art that various arrangements can be constructed although not specifically described or shown herein, but that these arrangements are within the scope of the present disclosure. Additionally, any example described herein is intended to help illustrate the principles of the present disclosure and is not intended to limit the scope of the disclosure to the specific arrangements described. Thus, the present disclosure is not limited to the specific embodiments or examples described below, but is instead limited by the claims and their equivalents.

[0041] The spars and spar parts described herein can be manufactured as composite parts. Composite parts, such as carbon fiber reinforced polymer (CFRP) parts, are initially laid up in multiple layers that are collectively referred to as a preform. The individual fibers within each layer of the preform are aligned parallel to one another, but different layers exhibit different fiber orientations in order to increase the strength of the resulting composite part along different dimensions. The preform includes a viscous resin that is cured in order to harden the preform into a composite part (e.g., for use in an aircraft). Carbon fibers that have been impregnated with uncured thermoset or thermoplastic resin are referred to as “prepregs.” Other types of carbon fibers include “dry fibers,” which have not yet been impregnated with thermoset resin, but can include a tackifier or adhesive. Dry fibers are infused with resin prior to hardening. For thermoset resins, hardening is a one-way process known as curing, whereas for thermoplastic resins, if the resin is reheated, the resin reaches a viscous form, after which the resin can be consolidated into a desired shape and cured. As used herein, the encompassing term for the process of converting a preform into a final hardened shape (i.e., converting a preform into a composite part) is referred to as “hardening,” and this term encompasses both the curing of thermoset preforms and the shaping / curing of thermoplastic preforms into a final desired shape.

[0042] Turning now to Figure 1This illustration depicts a representative aircraft 10, in which illustrative embodiments of wing panels and / or wing assemblies produced according to aspects of this disclosure can be implemented. In other words, aircraft 10 is an example of an aircraft that can be formed using composite parts, wing panels, and / or wing assemblies produced according to one or more aspects discussed herein. In this illustrative example, aircraft 10 has wings 12 attached to and extending toward either side of a fuselage 14. Aircraft 10 includes an engine 16 attached to each wing 12. A tail section 18 is disposed at the rear end of the fuselage 14, and the tail section 18 includes a vertical stabilizer 22 and a pair of opposing horizontal stabilizers 20. The wings 12 are formed by upper wing panels 30 and lower wing panels (not shown) joined together, wherein assemblies of ribs and spars (not shown) at least partially form the internal structure of the wing.

[0043] Figure 1A This is a block diagram of spar component 110 in an illustrative embodiment. As used herein, spar component 110 refers to a structural component. Furthermore, a spar component may be referred to as an aircraft spar component or a multi-segment spar component. Specifically, spar component 110 is a component that has been assembled with spar segments 112, 114, and 116 but has not yet been attached as part of a completed or nearly completed manufacturing process to other wing components (such as ribs or wing panels). After another component (bracket, reinforcement, fastener, etc.) is permanently attached to spar component 110, the spar component becomes either a rear spar 110-1 or a front spar 110-2. Figure 9 "Spar installation" (also referred to as "spar" in this document) refers to the rear spar 110-1 and front spar 110-2, which are attached as part of the wing box. Figure 9 The spar component 110 is one of the components that provides structural strength to the wing of the aircraft by extending from the inner end 197 to the outer end 199. For example, some spar components 110 extend from one side of the intersection point of the aircraft's body toward the wingtip 198. Some spar components 110 terminate at the wingtip 198, while in other embodiments, they terminate before reaching the wingtip 198. The spar component 110 can be considered as a spar (e.g., spar 110-1, 110-2) that provides most of the structural strength to the spar (e.g., spar 110-1, 110-2). Figure 9 The "core" of the wing spars 110-1 and 110-2. Figure 1A The diagram also illustrates multiple longitudinal portions 127 of the spar segments 112, 114, and 116 of the spar component 110. For clarity, spar segment 112 will sometimes be referred to as the first spar segment 112, spar segment 114 will sometimes be referred to as the second spar segment 114, and spar segment 116 will sometimes be referred to as the third spar segment 116.

[0044] In this embodiment, the spar component 110 consists of multiple spar segments 112, 114, 116 (see also...) Figure 3 The spar segments 112, 114, and 116 define segments / parts of the spar component 110 from the inner end 197 to the outer end 199. Each spar segment 112, 114, and 116 comprises multiple layers 122 consisting of fibers 124 (e.g., continuous carbon fibers, glass fibers, etc.) and resins 126 (e.g., thermosetting resins, thermoplastics, etc.). The spar component 110 is made from spar segments 112, 114, and 116 that are prefabricated and / or hardened prior to assembly. Therefore, the overall manufacturing process of the spars 110-1 and 110-2 manufactured using spar segments 112, 114, and 116 can be performed quickly and efficiently using equipment that is cheaper and less expensive than manufacturing systems that do not manufacture spars using multiple segments, for example, each spar segment 112, 114, and 116 in parallel.

[0045] Figure 2 Examples of illustrative embodiments are shown, such as Figure 1 The system 200 shown forms the spar component 110 using spar segments 112, 114, and 116. In this embodiment, the system 200 includes a layup machine 230, such as for dispensing fiber-reinforced material tows that form one or more single-layer sheets of flat charge 232 (e.g., planar filler) or fully flat charge 233 laminated on a mandrel 234 when combined with other flat charge 232. In this embodiment, the flat charge 232 and / or fully flat charge 233 are formed by a forming machine 240 to create a preform 242. The preform 242 is then laid onto a curing mandrel 238 by an end effector 262 of a pick-and-place (PNP) machine 260.

[0046] The preform 242 is a laminate of multiple flat plies 232 and / or a full flat ply 233 formed and assembled together. The one or more flat plies 232 are formed and laid up layer by layer (PBP) as part of the preform 242 by the forming machine 240. The shaping or forming of the flat plies 232 or full flat ply 233 is performed by the forming machine 240 by drape forming, stamp forming, layer by layer forming, or other suitable forming method. Manufacturing the preform 242 includes dispensing fiber reinforced material tows to form the flat plies 232, assembling the multiple flat plies 232 together into a laminate 235 to form a full flat ply 233, and shaping the full flat ply 233 into the preform 242 having a desired cross-sectional shape. As part of creating the preform 242, the forming machine 240 shapes the flat plies 232 to be laid up on a curing mandrel 238, which after application of a vacuum bag 252 and possibly a caul plate 253, the preform 242 is hardened into a spar segment 112, 114, 116 within an autoclave 250. Other alternatives have a combined vacuum bag and caul plate (not shown). The preform for the splice stiffener 340 Figure 3 shown in FIG. 3. The preform for the splice stiffener 340 can also be referred to as a splice stiffener preform.

[0047] In one embodiment not shown, the vacuum bag 252 covers the preforms 242 for multiple spar segments 112, 114, 116, and the entire assembly is co-cured. In this manner, the preforms 242 are assembled with the preform for the splice stiffener 340 and hardened into the spar part 110 integrated with the ribs 290 and the wing skin 30 in the autoclave 250. Figure 1 The preform for the splice stiffener 340 is assembled in a just in time (JIT) delivery manner to be assembled with the preforms 242.

[0048] As shown, spar segments 112, 114, 116 are hardened from preforms 242 at an autoclave 250. The plurality of spar segments 112, 114, 116 are then arranged end-to-end in alignment with one another. The plurality of spar segments 112, 114, 116 are combined with splice stiffeners 340 and assembled together with fasteners to form the spar part 110. The spar segments 112, 114, 116 are then assembled together via operations of a fastener installation station 272. In this embodiment, the fastener installation station 272 includes a fixture 276 that holds the spar segments 112, 114, 116 and the hardened splice stiffeners 340 relative to one another, and an end effector 274 that installs fasteners 278 through the components. That is, the end effector 274 drills installation holes and drives fasteners 278 (e.g., lock bolts) to join the splice stiffeners 340 and the spar segments 112, 114, 116. In this embodiment, the fasteners 278 are driven through the splice stiffeners 340 and the spar segments 112, 114, 116 to form the integral spar part 110. The hardened splice stiffeners 340 are assembled in a just-in-time production delivery manner for assembly with the plurality of spar segments 112, 114, 116.

[0049] The preforms 242 are hardened via the autoclave 250 into spar segments 112, 114, 116, respectively, and then arranged in end-to-end (in-line) alignment with one another. An end effector 271 lays down preforms for splice stiffeners 340 at splice regions 341, 341-1 between the spar segments 112, 114, 116. Figure 1A ) The end effector 271 lays down preforms for splice stiffeners 340 at splice regions 341, 341-1 between the spar segments 112, 114, 116. Then, at least in part, a press-clave 270 is used for a segment splicer 280 to co-bond the preforms for splice stiffeners 340 with the spar segments 112, 114, 116 in order to form the complete spar part 110. The press-clave 270 hardens one or more of the splice stiffeners 340 in place at the intersection points between the spar segments 112, 114, 116 via the application of heat and pressure.

[0050] The press-clave 270 process and the combination of the spar segments 112, 114, 116 with fastener 278 joining are also optional. The press-clave 270 process would precede the fastener installation process.

[0051] The operations of the system 200 are managed by a server 220 that includes a memory 222 and a controller 224. In one embodiment, the controller 224 is implemented as custom circuitry, a hardware processor executing programmed instructions stored in the memory 222, or some combination thereof.

[0052] The system 200 is capable of simultaneously manufacturing multiple sets of spar segments 112, 114, 116 for multiple spar parts 110. The system 200 is divided into several manufacturing spheres, each of which is capable of independently manufacturing a set of spar segments 112, 114, 116 that ultimately combine into a spar part 110. Thus, the hardening of the preforms 242 for the spar segments 112, 114, 116 is performed simultaneously with the assembly of the upstream preforms 242 from the flat fillers 232, the full flat fillers 233, and the forming machines 240. This results in the assembly of the spar segments 112, 114, 116 being performed at the same time (in parallel) as the hardening of the upstream set of spar segments 112, 114, 116 within the autoclave 250. The hardening splicing of yet another upstream set of spar segments 112, 114, and 116 by the segment splicer 280 into another spar part 110. The individual upstream and downstream spar segments 112, 114, 116 are generally used in sets for the front and rear spars 110-2 and 110-1 and / or right and left spars.

[0053] Figure 3is a drawing depicting a spar part 110 (having a width tapering from an inboard end 197 to an outboard end 199, and also having a length L) formed from a first spar segment 112, a second spar segment 114, and a third spar segment 116 in an illustrative embodiment. The first spar segment 112 includes a fiber reinforced material 301, the second spar segment 114 includes a fiber reinforced material 301-1 and is arranged in series with the first spar segment 112. An outboard end 312 of the first spar segment 112 is arranged end-to-end with an inboard end 321 of the second spar segment 114. The second spar segment 116 has an inboard end 321 and an outboard end 322. A splice reinforcement 340 covers a splice area 341 (also referred to as an "intersection" or "splice zone") including the outboard end 312 and the inboard end 321, and is placed in the splice area 341 between the first spar segment 112 and the second spar segment 114. According to embodiments, the splice area 341 can be implemented as a butt splice, scarf splice, lap splice, or other splice of the web and flanges of the outboard end 312 of the first spar segment 112 and the inboard end 321 of the second spar segment 114, with the splice reinforcement 340 completing the splice. In a similar manner, the third spar segment 116 includes a fiber reinforced material 301-2 and is arranged in series with the second spar segment 114. The third spar segment 116 has an inboard end 331 and an outboard end 332. The outboard end 322 of the second spar segment 114 is arranged end-to-end with the inboard end 331 of the third spar segment 116. Another splice reinforcement 340-1 covers a splice area 341-1 between the outboard end 322 of the second spar segment 114 and the inboard end 331 of the third spar segment 116. In one embodiment, an adhesive 319 (e.g., an epoxy, a thermoset resin, etc.) is used to bond the splice reinforcement 340, the first spar segment 112, and the second spar segment 114 into a unitary composite part. According to embodiments, the splice area 341 can be implemented as a butt splice, scarf splice, lap splice, or other splice of the web and flanges of the inboard end 331 of the third spar segment 116 and the outboard end 322 of the second spar segment 114, with the addition of the splice reinforcement 340-1 to complete the splice. In such embodiments, the adhesive 319 can also bond the other splice reinforcement 340-1, the second spar segment 114, and the third spar segment 116 together into the spar part 110. In further embodiments, the parts are fastened together, fastened and bonded together, co-cured or co-bonded together.

[0054] In Figure 3In particular embodiments, one of the spar segments (specifically, the second spar segment 114) entirely contains the kink 370, but in other embodiments, the kink 370 is distributed across the splice region 341, multiple kinks (not shown) each entirely contained within a spar segment 114, or each of multiple spar segments (such as the spar segment 112, the spar segment 114, and the spar segment 116) entirely contains the kink 370. Each kink 370 includes an inflection point 381 at which there is an intersection of the first neutral axis 350 and the second neutral axis 350-1 of the spar part 110. In one embodiment, this change is indicated by a kink angle Θ that is between 2 degrees and 10 degrees. In one illustrative example, the kink 370 includes a change in the axial direction of the spar part 110 that can be, for example, between 2 degrees and 10 degrees. In some embodiments, as shown, the kink 370 is separated 380 from the outboard neutral axis end point 322 (outboard end 322) and is separated from the inboard neutral axis end point (inboard end 321) of a spar segment (such as the second spar segment 114) by more than about 30 cm (1 foot). In another embodiment, the flat fillers 232 (or the resulting preform 242) in the Figure 3 Figure 2 In another embodiment, the flat fillers 232 (or the resulting preform 242) in the

[0055] Each of the spar segments 112, 114, and 116 assumes a different shape in order to account for the thickness of the kink 370 and / or the reduction in the kink angle Θ according to design parameters. Figure 3 The splice reinforcement 340 is further depicted to facilitate the integration of the spar segments 112, 114, and 116 into the spar part 110. A rib intersection 360 is disposed at and / or between the splice reinforcement 340, 340-1 and the receiving rib 290, which further supports the splice region 341, 341-1. The rib intersection 360 is shown as a small rectangular box, which is not intended to imply that the complex coupling of the rib 290 to the spar part 110 is limited to this relatively small area. In one embodiment, the rib intersection 360 is disposed on the spar part 110 opposite the splice reinforcement 340, as shown, on a side 391 opposite the side on which the splice reinforcement 340 is mounted. Thus, the splice reinforcement 340 is mounted on a side of the spar part 110 that is not visible from this view, while the rib 290 is mounted at the rib intersection 360 on a side 391 of the spar part 110 that is visible. The splice reinforcement 340 can be positioned across the rib intersection 360. ​

[0056] The spar component 110, assembled from individual spar segments 112, 114, and 116, offers substantial benefits over previous implementations because it reduces the size and complexity of the machines required to manufacture the spar component 110. Furthermore, the advantageous parallel manufacturing of spar segments 112, 114, and 116 significantly reduces manufacturing time and increases workload. This results in cost reduction and the technological benefit of saving factory floor space, which also increases workload. Based on the above discussion of the spar component 110 and the system for manufacturing the spar, the following... Figures 4-8 Various methods for manufacturing spar part 110 are described.

[0057] Regarding Figure 4 Illustrative details of the operation of system 200 are discussed. For this embodiment, it is assumed that the components of system 200 are awaiting activation in order to manufacture spar segments for assembly into one or more spars 110.

[0058] Figure 4 This is a flowchart illustrating a method 400 for manufacturing a spar component 110 via shared spar segments 112, 114, and 116 in an illustrative embodiment. (See also:) Figure 2 The steps of method 400 are described in system 200, but those skilled in the art will understand that method 400 can be performed in other systems. The steps in the flowchart described herein are not exhaustive and may include other steps not shown. The steps described herein may also be performed in an alternative order.

[0059] The layup machine 230, together with the molding machine 240, manufactures 402 preforms 242 of fiber-reinforced material (e.g., CFRP) for spar segments 112, 114, and 116. In one embodiment, manufacturing the 402 preform 242 includes the layup machine 230 laying one or more flat fillers 232. The flat fillers 232 are used directly to form the preform 242 or combined together to form a full flat filler 233, which is then used to form the preform 242. The flat fillers 232 and / or the full flat fillers 233 are then shaped by the molding machine 240 and laid down to form the preform 242 with a desired cross-section. In another embodiment, manufacturing the preform 242 includes laying fiber-reinforced material bundles onto a cured mandrel 238 that defines the shape of the preform 242. In yet another embodiment, manufacturing the preform 242 includes distributing fiber-reinforced material bundles that form the flat fillers 232, and forming the flat fillers 232 during the formation of the preform 242 with a desired cross-sectional shape.

[0060] Preform 242 is hardened 404 to form spar segments 112, 114, and 116. In one embodiment, the hardening 404 of preform 242 includes consolidation and solidification of a thermoplastic resin, while in a further embodiment, the hardening 404 includes heating a thermosetting resin to a curing temperature while under consolidation pressure in a pressurizer 250. In any embodiment, the resulting spar segments 112, 114, and 116 comprise hardened fiber-reinforced material shaped according to design parameters for a specific longitudinal portion 127 of spar part 110.

[0061] Spar segments 112, 114, and 116 are joined together 406 to form spar component 110, wherein outer ends 312 are butt-jointed or spliced ​​to inner ends 321, and outer ends 322 are butt-jointed or spliced ​​to inner ends 331. In one embodiment, this includes applying adhesive 319 or resin to spar segments 112, 114, and 116, and forming a splice (e.g., ...) between spar segments 110 in splice areas 341. Figure 3 (As shown). In another embodiment, this includes placing an unhardened “green” splice reinforcement 340 within the splice region 341 between spar segments 114 and 112, and hardening the splice reinforcement 340 by co-bonding in an extrusion vessel 270. This operation integrates the splice reinforcement 340 with the spar segments 114 and 112 to form a spliced ​​first spar segment 112 and second spar segment 114. Thus, in one embodiment, bonding the spar segments 112 and 114 together 406 includes applying the splice reinforcement 340, which partially covers the first spar segment 112 and second spar segment 114, and then bonding the splice reinforcement 340 406 to the first spar segment 112 and second spar segment 114. In some embodiments, the splice reinforcement 340 is positioned at a location where the splice reinforcement 340 extends across the rib intersection point 360 of the spar segment 110 to reinforce the splice. Therefore, in one embodiment, applying the splice reinforcement 340 includes clamping the first spar segment 112 and the second spar segment 114 between the front splice reinforcement 340 and the rear splice reinforcement (not shown). In these illustrative examples, the splice reinforcement 340 includes one of the front and rear splice reinforcements, the spar includes the other of the front and rear splice reinforcements, and the splice region 341 of the first spar segment 112 and the first splice region 341 of the second spar segment 114 are clamped between the front and rear splice reinforcements. In some illustrative examples, the splice reinforcement 340 clamps a portion of the first spar segment 112 and a portion of the second spar segment 114 between one of the second splice reinforcement or rib 290 and the splice reinforcement 340.

[0062] In another embodiment, applying the splice reinforcement 340 includes clamping the first spar segment 112 and the second spar segment 114 between the splice reinforcement 340 and the rib 290. In these illustrative examples, the splice region of the first spar segment and the first splice region of the second spar segment are clamped between the splice reinforcement and the rib. Similarly, the second spar segment 114 and the third spar segment 116 are spliced together using the splice reinforcement 340 and / or the rib 290, as discussed above, in a similar manner to the manner in which the first spar segment 112 is spliced to the second spar segment 114, resulting in the spar detail 110.

[0063] Upon completion, the spar detail 110 is assembled with additional details to form a complete spar, such as the aft spar 110-1 and the forward spar 110-2 of Figure 9 The method 400 provides substantial benefits over existing systems, as it enables the spar detail 110 to be formed from a series of smaller individual spar segments 112, 114, and 116 that each occupy a portion of the total length of the spar detail 110 through co-bonding. This reduces manufacturing complexity and time. Furthermore, if rework is warranted on an individual spar segment 112, 114, and 116, the rework can be performed on the individual spar segment 112, 114, and 116 rather than the entire spar detail 110.

[0064] Figure 5is a flowchart showing another method 500 of fabricating spar part 110 with spar segments 112, 114, and 116 via co-curing in illustrative embodiments. According to method 500, preforms 242 of fiber-reinforced material for spar segments 112, 114, and 116 are fabricated 502, and can be accomplished in a similar manner as the fabrication 402 steps described above. For example, prior to co-curing, first spar segment 112, second spar segment 114, third spar segment 116, splice reinforcement 340, and second splice reinforcement 340-1 can comprise dry fiber infused in place with resin prior to co-curing. As part of this fabrication process, preforms 242 can comprise scarf-type splicing of complementary features of preforms 242 within splice region 341. Outboard end 312 is butted or spliced to inboard end 321 within splice region 341. Outboard end 322 is butted or spliced to inboard end 331 within splice region 341-1. In one embodiment, fabricating 502 preforms 242 comprises dispensing / laying fiber-reinforced material tows to form flat packings 232, and shaping flat packings 232 into preforms 242 having a desired cross-sectional shape. In another embodiment, laying preforms 242 comprises combining a plurality of flat packings 232 together into a full flat packing 233, and shaping full flat packing 233 into preform 242 having a desired cross-sectional shape.

[0065] Preforms 242 for spar segments 112, 114 are spliced 504 together with outboard end 312 of first spar segment 112, which is arranged end-to-end with inboard spar segment 114. In other words, spar segments 112, 114 are arranged in series. In one embodiment, splicing 504 preforms 242 for spar segments 112, 114 is performed by lap splicing, butt splicing, and / or scarf splicing of complementary slopes or patterns of preforms within splice region 341. Preforms 242 for spar segments 114, 116 are spliced 504 together with outboard end 322 of first spar segment 112, which is arranged end-to-end with inboard end 331 of spar segment 116. In one embodiment, splicing 504 preforms 242 for spar segments 114, 116 is performed by lap splicing, butt splicing, and / or scarf splicing of complementary slopes or patterns of preforms 242 within splice region 341-1. In further embodiments, splicing 504 preforms 242 for spar segments 114, 116 is performed by arranging dry fiber preforms together and then infusing preforms 242 with resin.

[0066] Next, the PNP machine 260 applies 506 a preform for a splice reinforcement 340 to the splice area between the preforms 242 for the spar segments 112, 114. In one embodiment, this operation includes laying up an unhardened preform for the splice reinforcement 340 in the splice area 341 that forms a splice between the preforms 242 for the spar segments 112, 114. In another embodiment, applying 506 the preform for the splice reinforcement 340 includes picking up and laying up the preform for the splice reinforcement 340 at a location where the preform for the splice reinforcement 340 extends across the rib intersection point 360 of the spar part 110. In another embodiment, applying 506 the preform for the splice reinforcement 340 includes laying up the preform for the splice reinforcement 340 (not shown) at a location where the preform for the splice reinforcement 340 extends across the rib intersection point 360 of the spar part 110. In one embodiment, applying 506 the preform for the splice reinforcement 340 (not shown) includes laying up the preform for the splice reinforcement 340 across the outboard end 312 and the inboard end 321 and across a side of the spar part 110 opposite the rib intersection point 360 of the spar part 110. In this manner, the rib intersection point 360 is disposed on the spar part 110 opposite the splice reinforcement 340. Further, applying the preform for the splice reinforcement 340-1 (not shown) includes laying up the preform for the splice reinforcement 340-1 across the outboard end 312 and the inboard end 321 and placed on a side of the spar part 110 opposite the rib intersection point 360 for the spar part 110. Similarly, the rib intersection point 360 is disposed on the spar part 110 opposite the splice reinforcement 340-1.

[0067] While the preforms 242 for the spar segments 112, 114, and 116 are being spliced 504 together, the autoclave 250 cures 508 the preforms 242 for the spar segments 112, 114, and the preform for the splice stiffener 340, 340-1 to form the complete spar part 110. In further embodiments, the curing 508 of the preforms 242 for the spar segments 112, 114, and 116 and the preform for the splice stiffener 340, 340-1 includes vacuum bagging the preforms 242 for the spar segments 112, 114, and 116 and the preform for the splice stiffener 340, 340-1. The preforms 242 for the spar segments 112, 114, and 116 and the preform for the splice stiffener 340, 340-1 are consolidated via vacuum compaction and the preforms 242 for the spar segments 112, 114, and 116 and the preform for the splice stiffener 340 are heated to a cure or consolidation temperature. Thus, in one embodiment, the vacuum bag 252 additionally covers the preform 242 for the third spar segment 116 and the preform for the second splice stiffener 340, 340-1.

[0068] Figure 6 is a flowchart illustrating another method 600 of manufacturing a spar part 110 from spar segments 112, 114, and 116 via fasteners in an illustrative embodiment. A layup machine 230 and / or a forming machine 240 manufactures 602 a preform 242 of fiber-reinforced material for the spar segments 112, 114, and 116. The manufacture 602 can be performed in a similar manner as the manufacture 402 of the preform 242 discussed above. Figure 4

[0069] ​The preforms 242 are hardened 604 to form the spar segments 112, 114, and 116. In one embodiment, this includes operating the autoclave 250 to harden 604 the preforms 242 into individual spar segments 112, 114, and 116. At this point, the spar segments 112, 114, and 116 remain physically separate from one another. In another embodiment, the splice stiffeners 340, 340-1 are utilized and fastened to each of the spar segments 112, 114, and 116 to form a splice.

[0070] The clamps 276 of the fastener installation station 272 hold the spar segments 112 and 114 and the additional spar segments 114 and 116 and the hardened splice stiffeners 340, 340-1 in the splice areas 341, 341-1, respectively. The end effector 274 installs fasteners to join these components into the spar part 110. Specifically, the end effector 274 applies 606 fasteners to couple the spar segment 112 and the spar segment 114 and the spar segment 114 and the spar segment 116 together to form the complete spar part 110. In one embodiment, applying 606 the fasteners includes applying the splice stiffeners 340, 340-1 that partially cover the outboard end 312 and the inboard end 321 and installing fasteners through the splice stiffeners 340 and through the first spar segment 112 and the second spar segment 114. The fasteners are applied 606 to the splice stiffeners 340-1 that partially cover the outboard end 322 and the inboard end 331 and installing fasteners through the splice stiffeners 340-1 and through the first spar segment 112 and the second spar segment 114. In another embodiment, applying the splice stiffeners 340 includes placing the splice stiffeners 340 at locations where the splice stiffeners 340 extend across the rib intersection points 360 for the spar part 110 located in the splice areas 341, 341-1. In one embodiment, applying 606 the fasteners for the splice stiffeners 340 includes placing the splice stiffeners 340 across the outboard end 312 and the inboard end 321 and across a side of the spar part 110 opposite the rib intersection points 360 for the spar part 110. In this way, the rib intersection points 360 are disposed on the spar part 110 opposite the splice stiffeners 340. After assembly, the spar part 110 is combined with other parts into the spar 110-1, 110-2 Figure 9 ) in this way, the fastener installation station 272 assembles the splice stiffeners 340, 340-1 and the spar segments 112, 114, and 116 into a single integral spar part 110.

[0071] Figure 7is a flowchart showing a method 700 for manufacturing a spar part 110 having one or more kinks 370 in illustrative embodiments. The method 700 includes laying up or manufacturing 702 a preform 242 of fiber-reinforced material that will be assembled into a spar part 110 including one or more kinks 370. Each kink 370 is entirely contained within a single preform 242 for a spar segment 112, 114, and 116, e.g., spar segments 112, 114, 116. In one embodiment, manufacturing 702 the preform 242 includes changing a direction from a first neutral axis 350 to a second neutral axis 350-1 at each kink 370 (e.g., during layup or forming), where the change is between 2 and 10 degrees. In further embodiments, manufacturing 702 the preform 242 includes placing each kink 370 at least 30 cm (1 foot) from an outboard end 322 and an inboard end 321 of the preform 242 for the spar segment 114. Likewise, it is also contemplated to place kinks 370 in the spar segment 112 and / or the spar segment 114. In yet another embodiment, laying up (i.e., manufacturing 702) the preform 242 includes laying up a flat charge 232 on a mandrel 234 and then shaping the flat charge 232 on a rigid tool such as a cure mandrel 238. The shaping or forming of the flat charge 232 or full flat charge 233 is performed by drape forming, stamp forming, layer-by-layer forming, or other suitable forming methods prior to incorporation into the preform 242.

[0072] The preform 242 is hardened 704 to form the spar segments 112, 114, and 116, which are then assembled 706 together to form the complete spar part 110 exhibiting the kinks 370. As described above, the assembly process can be performed via co-curing, co-bonding, or via fasteners or a combination of fasteners and bonding or co-bonding. In one embodiment, the hardening 704 of the preforms 242 for the spar segments 112, 114, and 116 is performed at the same time as the assembly of the upstream preforms 242 with the flat charges 232, full flat charges 233, and forming machines 240 together, such that the assembly of the spar segments 112, 114, and 116 is performed at the same time as the hardening 704 of a group of upstream spar segments 112, 114, and 116 within the autoclave 250 and the post-hardening splicing of yet another group of upstream spar segments 112, 114, and 116 into the spar part 110 by the segment splicer 280. The upstream spar segments and the downstream spar segments are typically grouped for the front spar 110-2 and the rear spar 110-1 and / or the right spar and the left spar.

[0073] Assembling 706 together spar segments 112 and 114 to form a portion of complete spar part 110 exhibiting kink 370. In one embodiment, assembling 706 spar segments 112 and 114 includes applying splice reinforcement 340 across outboard end 312 and inboard end 321 to splice spar segment 112 to spar segment 114. Assembling 706 spar segments 114 and 116 includes applying splice reinforcement 340-1 across outboard end 322 and inboard end 331 to splice spar segment 114 to spar segment 116. In some illustrative examples, assembling 706 spar segments 112, 114, 116 includes forming at least one of a lap splice, a butt splice, and a scarf splice between adjacent ends of spar segments 112, 114, 116, and attaching splice reinforcement 340, 340-1 to the splice region 341, 341-1 using one of co-curing, co-bonding, and fasteners. In some illustrative examples, assembling 706 spar segments 112, 114, 116 includes arranging spar segments 112, 114, 116 in series alignment with each other to form splice region 341, 341-1, and applying splice reinforcement 340, 340-1 to spar segments 112, 114, 116 within splice region 341, 341-1.

[0074] While a single kink 370 can be implemented in some applications, in other applications, a single kink 370 can result in an angle between the two neutral axes that is too large to manufacture with known CFRP manufacturing methods. To this end, Figure 8 is a flowchart illustrating another method 800 for manufacturing spar part 110. In the embodiment illustrated in method 800, multiple kinks are joined in close proximity to each other across adjacent spar segments (e.g., spar segments 112 and 114) to more smoothly transition the change in neutral axis between adjacent spar segments 112, 114. In such embodiments, for clarity, the multiple kinks are referred to herein as sub-kinks 371, 371-1 and are illustrated in Figure 13 .

[0075] In some illustrative examples, manufacturing 702 preform 242 includes changing an axial direction of preform 242 at each kink 370. In some illustrative examples, manufacturing 702 preform 242 includes changing an axial direction of preform 242 at each kink 370, the change in axial direction being a kink angle between 2 degrees and 10 degrees. In some illustrative examples, each preform 242 includes a first end and a second end opposite the first end, and manufacturing 702 preform 242 includes placing each kink 370 at least one foot from an end of preform 242.

[0076] In accordance with Figure 8The preforms 242 of fiber-reinforced material are manufactured 802 as spar segments 112, 114, 116 that will be assembled into the spar part 110 including multiple sub-kinks 371, 371-1. For example, adjoining ends of the spar segment 112 and the spar segment 114 can each include a sub-kink, such as the sub-kinks 371, 371-1, and each sub-kink 371, 371-1 changes the orientation from a first neutral axis 350 to a second neutral axis 350-1 of the preforms 242 of the spar part 110. Manufacturing 802 the preforms 242 can be performed as described above for the previously described method.

[0077] The ends of the preforms 242 having sub-kinks 371, 371-1 (e.g., the spar segments 112, 114) are aligned 804 so that the sub-kinks 371, 371-1 are proximate to each other within the splice region, which will be described further with reference to Figure 13 The spar segments 112 and 114 are then joined 806 together within the splice region. Aligning 804 and joining 806 the ends includes overlapping, butt splicing, scarf splicing, or other means during formation of the preforms 242.

[0078] The spar part 110 is manufactured using the preforms 242. In one embodiment, manufacturing the spar part 110 is performed by co-curing the preforms 242, co-bonding the spar segments 112, 114, and 116, fastening the spar segments 112, 114, and 116, or co-bonding and fastening the spar segments 112, 114, and 116 together, as discussed above.

[0079] Figure 9 An aircraft wing 900 including a spar is depicted in an illustrative embodiment. Figure 9A A cross-section of the wing 900 (similar to the wing 12 in Figure 9 corresponds to the view arrow 9A of the wing 900 (similar to the wing 12 in Figure 1 In an illustrative embodiment, the wing 900 is of an illustrative embodiment of an aircraft including an aft spar 110-1 and a forward spar 110-2. The aft spar 110-1 and the forward spar 110-2 are hidden under a wing panel 75 (similar to the panel 30 in Figure 1 The aft spar 110-1 and the forward spar 110-2 terminate at the wing tip 198. As in Figure 9AAs shown, the rear spar 110-1 includes a first flange 902 and a second flange 906 with a web 904 therebetween. For simplicity, the front spar 110-2 is also shown with a first flange 902, a second flange 906, and a web 904.

[0080] Typical wing panels 75, 76 include a composite wing skin and a composite stringer attached to the wing skin and extending from an inboard end 197 to an outboard end 199. As will be discussed below, the wing panels 75, 76 are formed from composite materials, such as CFRP laid up in tape widths with various fiber orientations to provide the desired strength and flexibility. As Figure 9 As shown, the rib 290 is disposed between the rear spar 110-1 and the front spar 110-2 below the wing panel 75 in the chordwise direction.

[0081] Figure 9A The front spar 110-2 and the rear spar 110-1 are depicted, which are disposed between the wing panels 75, 76 and joined to the rib 290. The wing panel 75 is coupled to the flange 906 and the rib 290, respectively, while the wing panel 76 is coupled to the flange 902 and the rib 290, respectively, to complete the wing 900. For clarity, the wing panels 75 and 76 are not shown with stringers.

[0082] Figures 10-11 A spar part 110 in an illustrative embodiment is depicted. It represents either the rear spar 110-1 and / or the front spar 110-2. Figure 10 A first view is shown, in which the web 904 and the flange 902 are visible, while Figure 11 A second view is shown corresponding to the view arrow 11 of Figure 10 In the Figure 11 In the Figure 12 In the Figure 12 In the

[0083] Figure 12 A kink 370 at a portion of the spar part 110 in an illustrative embodiment is depicted, and corresponds to the label Figure 3 of Figure 12 of the box. The kink 370 is a result of the first neutral axis 350 of the spar part 110 (shown as a dashed line) bending at the inflection point 381 by an angle Θ, resulting in the second neutral axis 350-1. The kink 370 is integrated into the spar part 110 by a plurality of segments Figure 12The formation of a spar component 110 (not shown) could complicate the assembly process. Therefore, in one embodiment, the bend 370 is entirely contained within a single spar segment 112, 114, and 116. Figure 12 As shown, bend 370 is completely contained within the second spar segment 114. Because bend 370 is completely contained within spar segments 112, 114, and 116, the different spar segments 112, 114, and 116 in spar component 110 are in splicing areas 341 and / or 341-1 (all within... Figure 3 As shown in the diagram, splicing areas 341 and / or 341-1 exhibit neutral axes aligned with the first neutral axis 350 and / or the second neutral axis 350-1, respectively. This alignment with the first neutral axis 350 and / or the second neutral axis 350-1 substantially reduces the difficulty of aligning and securing the spar segments 112, 114, and 116 together.

[0084] In a further embodiment, such as in Figure 13 and Figure 14 As shown, the bend 370 with angle θ is subdivided into two sub-bends 371 and 371-1 at bends 1312 and 1322, respectively, with angular offsets of θ2 and θ3 at adjacent spar segments (such as spar segments 112 and 114). At bend 1322, the first neutral axis 1324 transforms into a second neutral axis 242-1 at angle θ2. Sub-bend 371-1 is the result of bending the first neutral axis 1324 of spar component 110 at bend point 381-1 by angle θ2, thus obtaining the second neutral axis 242-1. At bend 1312, the second neutral axis 242-1 transforms into a third neutral axis 1314 at angle θ3. Sub-bend 371 is the result of bending the second neutral axis 242-1 of the spar component 110 at bend point 381-2 at an angle θ3, thereby obtaining the third neutral axis 1314. In an embodiment, the sum of the angle θ2 of sub-bend 371-1 and the angle θ3 of sub-bend 371 equals the angle θ, which may or may not be equal to... Figure 12 The angle θ of the single bend 370 configuration is described in the text.

[0085] In another embodiment, sub-turns 371, 371-1 are distributed on the additional spar segments 112, 114, 116. For example, the intersection of spar segments 112 and 114 may include two sub-turns 371, 371-1, and the intersection of spar segments 114 and 116 may also include two sub-turns 371, 371-1.

[0086] Therefore, when the spar segments 112, 114, and 116 are placed together, they can be aligned together by aligning them (804) and applying (Figure 3 splice reinforcement 340 to secure the joint 806 for accurate execution of the assembly process. The splice reinforcement 340 can be co-cured, co-bonded, or fastened to both of the corresponding spar segments 112, 114, 116.

[0087] To reiterate, in further embodiments, a first spar segment (e.g., spar segment 112) includes fiber-reinforced material and includes a sub-kink 371-1 disposed at an end 1310, and a second spar segment (e.g., spar segment 114) also includes fiber-reinforced material, including another sub-kink 371 disposed at an end 1320. In some illustrative examples, the sub-kink 371-1 associated with the first spar segment 112 is disposed within about 30 cm (1 foot) of an end of the first spar segment 112; and the sub-kink 371 associated with the second spar segment 114 is disposed within about 30 cm (1 foot) of an end of the second spar segment 114. The end 1310 is arranged in-line and adjacent (e.g., within 30 cm / 1 foot of) the end 1320, and the spar segments 112, 114 of this example are arranged in-line such that the sub-kinks 371, 371-1 together provide a desired amount of kink. The components (e.g., splice reinforcement 340 as described above) Figure 3 The splice reinforcement 340 can structurally integrate the first spar segment 112 and the second spar segment 114. In some illustrative examples, the splice reinforcement 340 is arranged to sandwich the splice region 341 and the sub-kinks 371, 371-1 between the splice reinforcement 340 and a second splice reinforcement. In some illustrative examples, the manufacturing 802 the preform 242 includes manufacturing 802 the preform 242 such that each sub-kink 371-1, 371 has an equal angular offset. In some illustrative examples, the manufacturing 802 the preform 242 includes manufacturing 802 the preform 242 such that the sub-kinks 371-1, 371 have unequal angular offsets. In some illustrative examples, each sub-kink 371-1, 371 changes the axial direction of the spar part 110 by half of the amount of the kink 370.

[0088] Figure 14 Embodiments are depicted in which a spar segment (e.g., spar segment 116) presents one bend 1312 that differs in size from another bend 1322 that can be found in the spar segment 114. The sum of the angle Θ2 of the sub-kink 371-1 plus the angle Θ3 of the sub-kink 371 still equals the angle Θ, but the angle Θ2 is not equal to the angle Θ3.

[0089] In some embodiments, sub-corners 371, 371-1 are formed at the bends 1312 and 1322 of the flat filler 232 that have bends 1312 and 1322 when laid up. Another alternative has the forming machine shape the flat filler 232 without sub-corners 371-1, 371-1 into the preform 242 that includes sub-corners 371, 371-1 and bends 1312 and 1322 at the end portions 1310 and 1320 that combine to form the corner of angle Θ.

[0090] Figures 15-17 Depicted is the forming of the flat filler 232 into the preform 242 having the predetermined cross-section 1702 in an illustrative embodiment. In Figure 15 , one or more layers of the flat filler 232 are formed by applying a plurality of tows 1510. In Figure 15 , the view arrow 16 corresponds to the view arrow 16 in Figure 16 , the full flat filler including a plurality of flat fillers 232 is transferred to the mandrel 234 having the profile 1610. In Figure 17 , the flat fillers 232 are shaped to conform to the profile 1610 by applying pressure and / or heat, resulting in the preform 242. Although the full flat filler 233 having a plurality of flat fillers 232 is depicted as being shaped to conform to the profile 1610, in some un-depicted examples, each flat filler 232 is shaped individually to conform to the profile 1610.

[0091] Figure 18 Depicted is a scarf joint 1850 (or other overlap) between spar segments 112 and 114 in an illustrative embodiment. As Figure 18 indicated, the spar segment 112 includes an end portion 1834 that overlaps an end portion 1844 of the spar segment 114. The end portions overlap along ramps 1832 and 1842 of the spar segments 112, 114. The ramp rates of each ramp 1832, 1842 can be complementary and / or equal such that a uniform thickness is maintained across the scarf joint 1850. For clarity, the end portion 1834 and the end portion 1844 are simply depicted without the flanges 902, 906. In another embodiment, the scarf joint 1850 is one of a plurality of scarf joints that join a plurality of spar segments together into a single integral spar part. The scarf joint 1850 can be fabricated by co-curing, co-bonding, fastener installation, or other means as desired. Instead of the scarf joint 1850, a stepped lap joint or other suitable joint can be envisioned.

[0092] Examples

[0093] In the following examples, additional processes, systems, and methods are described in the context of manufacturing a spar part.

[0094] With more particular reference to the drawings, embodiments of the disclosure can be described in the context of aircraft manufacturing and servicing in the method 1900 shown Figure 19 and in the context of an aircraft 1902 shown in Figure 1 . Figure 19 is a flow diagram of an aircraft operation and servicing method in illustrative embodiments. During pre-production, the method 1900 can include specification and design 1904 of the aircraft 1902 and material procurement 1906. During production, component and subassembly manufacturing 1908 and system integration 1910 of the aircraft 1902 takes place. Thereafter, the aircraft 1902 can go through certification and delivery 1912 in order to be placed in service 1914. While in service by a customer, the aircraft 1902 is scheduled for routine work including maintenance and service 1916, which can also include modification, refurbishment, refit, and / or the like. The apparatus and methods embodied herein can be employed during any one or more suitable stages of production and servicing in the method 1900 (e.g., specification and design 1904, material procurement 1906, component and subassembly manufacturing 1908, system integration 1910, certification and delivery 1912, placement in service 1914, maintenance and service 1916) and / or during any one or more suitable stages of any suitable component of the aircraft 1902 (e.g., airframe 1918, systems 1920, interior 1922, propulsion system 1924, electrical system 1926, hydraulic system 1928, environmental system 1930).

[0095] Each of the processes of method 1900 can be performed or carried out by a system integrator, a third party, and / or an operator (e.g., a customer). For the purposes of this description, a system integrator can include without limitation any number of aircraft manufacturers and major-system subcontractors; a third party can include without limitation any number of vendors, subcontractors, and suppliers; and an operator can be an airline, leasee, military entity, service organization, and / or other operator.

[0096] Figure 20 is a block diagram of an aircraft in illustrative embodiments. As shown in Figure 20 , the aircraft 1902 produced by the method 1900 can include an airframe 1918 with a plurality of systems 1920 and an interior 1922. Examples of systems 1920 include one or more of a propulsion system 1924, an electrical system 1926, a hydraulic system 1928, and an environmental system 1930. Any number of other systems can be included. Although an aerospace example is shown, the principles of the application can be applied to other industries, such as the automotive industry.

[0097] As already mentioned above, the apparatus and methods embodied herein can be employed during any one or more of the stages of production and maintenance described in method 1900. For example, components or subassemblies corresponding to component and subassembly manufacturing 1908 can be fabricated or manufactured similar to that of the produced components or subassemblies for aircraft 1902 while the aircraft 1902 is in service, e.g. in flight, on the ground, and / or in a maintenance depot. Moreover, one or more apparatus embodiments, method embodiments or a combination thereof can be employed during the subassembly manufacturing 1908 and system integration 1910 stages, e.g. to significantly reduce the assembly time of aircraft 1902 or to reduce the cost of aircraft 1902. Similarly, one or more of the apparatus embodiments, method embodiments or a combination thereof can be utilized during the in-service 1914 and maintenance and service 1916 stages, e.g. to refurbish, repair or otherwise maintain and / or service aircraft 1902. Thus, the present application can be employed at any stage of aircraft 1902, or any combination thereof, such as specification and design 1904, materials procurement 1906, component and subassembly manufacturing 1908, system integration 1910, certification and delivery 1912, in-service 1914, maintenance and service 1916, and / or any suitable component of aircraft 1902 (e.g. fuselage 1918, systems 1920, interior 1922, propulsion system 1924, electrical system 1926, hydraulic system 1928, and / or environmental system 1930).

[0098] In one embodiment, the part comprises a portion of the fuselage 1918 and is manufactured during component and subassembly manufacturing 1908. The part can then be assembled into an aircraft during system integration 1910 and then used in-service 1914 until wear makes the part unusable. The part can then be discarded and replaced with a newly manufactured part during maintenance and service 1916. The components and methods of the present application can be used throughout component and subassembly manufacturing 1908 in order to manufacture new parts.

[0099] Further, the present disclosure includes examples in accordance with the following clauses.

[0100] Clause 1. A method 400 for manufacturing a spar part 110 for an aircraft 10, the method 400 comprising: manufacturing 402 a preform 242 of fiber-reinforced material 301 for a spar segment 112, 114, 116; hardening 404 the preform 242 to form the spar segment 112, 114, 116; and joining together the spar segments 112, 114, 116 to form the complete spar part 110.

[0101] Clause 2. The method 400 of clause 1, wherein joining the spar segments 112, 114, 116 together comprises: applying splice reinforcements 340 that partially cover the first spar segment 112 and the second spar segment 114; and bonding 406 the splice reinforcements 340 to the first spar segment 112 and the second spar segment 114.

[0102] Clause 3. The method 400 of clause 2, wherein applying the splice reinforcements 340 comprises clamping the splice reinforcements 340 between the splice areas 341 of the first spar segment 112 and the second spar segment 114.

[0103] Clause 4. The method 400 of clause 2 or 3, wherein applying the splice reinforcements 340 comprises clamping the splice areas 341 of the first spar segment 112 and the second spar segment 114 between a front splice reinforcement and a back splice reinforcement.

[0104] Clause 5. The method 400 of clause 2, 3, or 4, wherein applying the splice reinforcements 340 comprises placing the splice reinforcements 340 at locations where the splice reinforcements 340 extend across rib intersection points 360 of the spar part 110.

[0105] Clause 6. The method 400 of any of clauses 2-5, wherein applying the splice reinforcements 340 comprises clamping the first spar segment 112 and the second spar segment 114 between the splice reinforcements 340 and the ribs 290.

[0106] Clause 7. The method 400 of any of clauses 2-6, wherein bonding 406 the splice reinforcements 340 comprises hardening the splice reinforcements 340 by co-bonding in an extrusion kettle.

[0107] Clause 8. The method 400 of any of the preceding clauses, wherein joining 406 the spar segments 112, 114, 116 together comprises: applying one of an adhesive and a resin to the spar segments 112, 114, 116; and forming a splice between the spar segments 112, 114, 116 in the splice areas 341 of the spar segments 112, 114, 116.

[0108] Clause 9. The method 400 of any of the preceding clauses, wherein joining 406 the spar segments 112, 114, 116 together comprises forming one of a butt splice, a lap splice, and a scarf splice between the splice areas 341 of the first spar segment 112 and the second spar segment 114.

[0109] Clause 10. A spar for an aircraft 10, the spar comprising: a first spar segment 112 comprising a fiber reinforced material 301, the first spar segment 112 comprising a splice region 341; a second spar segment 114 comprising a fiber reinforced material 301-1, the second spar segment 114 comprising a first splice region 341 and disposed in series with the first spar segment 112; and a splice reinforcement 340 covering at least a portion of the first splice region 341 of the second spar segment 114 and the splice region 341 of the first spar segment 112, the splice reinforcement 340 bonded to the first spar segment 112 and the second spar segment 114.

[0110] Clause 11. The spar of Clause 10, wherein the spar is one of a front spar 110-1 and a rear spar 110-2.

[0111] Clause 12. The spar of Clause 10 or 11, wherein the splice reinforcement 340 is sandwiched between the splice region 341 of the first spar segment 112 and the first splice region 341 of the second spar segment 114.

[0112] Clause 13. The spar of Clause 10, 11, or 12, wherein the splice reinforcement 340 comprises a front splice reinforcement and a rear splice reinforcement, the splice region 341 of the first spar segment 112 and the first splice region 341 of the second spar segment 114 being sandwiched between the front splice reinforcement and the rear splice reinforcement.

[0113] Clause 14. The spar of any of Clauses 10-13, wherein the splice reinforcement 340 extends across a rib intersection point 360 of the spar.

[0114] Clause 15. The spar of any of Clauses 10-14, wherein the splice region 341 of the first spar segment 112 and the first splice region 341 of the second spar segment 114 are sandwiched between the splice reinforcement 340 and a rib.

[0115] Clause 16. The spar of any of Clauses 10-15, wherein the splice reinforcement 340 is co-bonded to the splice region 341 of the first spar segment 112 and the first splice region 341 of the second spar segment 114.

[0116] Clause 17. The spar of any of clauses 10-16, wherein the second spar segment 114 further comprises a second splice region 341-1, the spar further comprising: a third spar segment 116 comprising a fiber-reinforced material 301-2, the third spar segment 116 comprising the splice region 341-1 and being disposed in series with the second spar segment 114; and a second splice reinforcement 340-1 covering at least a portion of the splice region 341-1 of the third spar segment 116 and at least a portion of the second splice region 341-1 of the second spar segment 114, the second splice reinforcement 340-1 being bonded to the second spar segment 114 and the third spar segment 116.

[0117] Clause 18. Manufacturing a portion of an aircraft 10 using the spar of any of clauses 10-17.

[0118] Clause 19. An aircraft wing 12 comprising a multi-segment spar part 110, the spar part 110 comprising: a first spar segment 112 comprising a fiber-reinforced material 301, the first spar segment 112 comprising a splice region 341; a second spar segment 114 comprising a fiber-reinforced material 301-1, the second spar segment 114 comprising a first splice region 341, the second spar segment 114 being disposed in series with the first spar segment 112; a splice reinforcement 340 covering at least a portion of the first splice region 341 of the second spar segment 114 and the splice region 341 of the first spar segment 112, the splice reinforcement 340 being co-bonded to the first spar segment 112 and the second spar segment 114; a third spar segment 116 comprising a fiber-reinforced material 301-2, the third spar segment 116 comprising a splice region 341-1 and being disposed in series with the second spar segment 114; and a second splice reinforcement 340-1 covering at least a portion of the splice region 341-1 of the third spar segment 116 and at least a portion of the second splice region 341-1 of the second spar segment 114, the second splice reinforcement 340-1 being co-bonded to the second spar segment 114 and the third spar segment 116.

[0119] Clause 20. The aircraft wing 12 of clause 19, wherein at least one of the splice reinforcement 340 and the second splice reinforcement 340-1 extends across a rib intersection 360 of the multi-segment spar part 110.

[0120] Clause 21. A method 500 for manufacturing a spar part 110 for an aircraft 10, the method 500 comprising: manufacturing 502 a preform 242 of fiber-reinforced material 301, 301-1 for a first spar segment 112 and a second spar segment 114; splicing 504 an end 312 of the preform 242 of the first spar segment 112 to an end 321 of the preform 242 of the second spar segment 114 to define a splice region 341; applying 506 at least one preform for a splice stiffener 340 to the splice region 341; and simultaneously hardening 508 the preforms for the spar segments 112, 114 and the splice stiffener 340 to form a portion of the spar part 110.

[0121] Clause 22. The method 500 of clause 21, wherein simultaneously hardening 508 the preforms 242 for the spar segments 112, 114 and the preform for the splice stiffener 340 comprises: vacuum bagging the preforms 242 for the spar segments 112, 114 and the preform for the splice stiffener 340 with a vacuum bag 252; consolidating the preforms 242 for the spar segments 112, 114 and the preform for the splice stiffener 340 via the vacuum bag 252; and heating the preforms 242 for the spar segments 112, 114 and the preform for the splice stiffener 340.

[0122] Clause 23. The method 500 of clause 21 or 22, wherein applying 506 the at least one preform for the splice stiffener 340 comprises picking up and laying down the preform for the splice stiffener 340 such that the splice stiffener 340 extends across a rib intersection 360 of the spar part 110.

[0123] Clause 24. The method 500 of clause 21, 22, or 23, wherein applying 506 the at least one preform for the splice stiffener 340 comprises laying the preform for the splice stiffener 340 onto the spliced ends 312, 321 of the first spar segment 112 and the second spar segment 114 such that the splice stiffener 340 extends across a rib intersection 360 of the spar part 110.

[0124] Clause 25. The method 500 of any of clauses 21 to 24, wherein applying 506 the at least one preform for the splice stiffener 340 comprises laying the preform for the splice stiffener 340 across a side of the first spar segment 112 and the second spar segment 114 opposite a rib intersection 360 of the spar part 110.

[0125] Clause 26. The method 500 of any of Clauses 21-25, wherein simultaneously hardening 508 the preforms 242 for the spar segments 112, 114 and the splice reinforcement 340 comprises: arranging dry fiber preforms for the first spar segment 112, the second spar segment 114 and the splice reinforcement 340; and infusing the dry fibers with resin.

[0126] Clause 27. The method 500 of any of Clauses 21-26, wherein splicing 508 the end 312 of the first spar segment 112 preform 242 to the end 321 of the second spar segment 114 preform 242 to define a splice region 341 comprises splicing the ends 312, 321 with at least one of a lap splice, a butt splice and a scarf splice.

[0127] Clause 28. The method 500 of any of Clauses 21-27, wherein manufacturing 502 the preforms 242 comprises: dispensing fiber reinforcement 301 tows to form flat packings 232; and shaping the flat packings 232 into the preforms 242 having a desired cross-sectional shape.

[0128] Clause 29. The method 500 of any of Clauses 21-28, wherein manufacturing 502 the preforms 242 comprises: dispensing fiber reinforcement 301 tows to form flat packings 232; combining a plurality of the flat packings 232 together into a full flat packing 233; and shaping the full flat packing 233 into the preform 242 having a desired cross-sectional shape.

[0129] Clause 30. A portion of an aircraft 10 assembled according to the method of any of Clauses 21-29.

[0130] Clause 31. A spar part 110 comprising: a preform 242 for a first spar segment 112, the first spar segment 112 comprising a splice region 341; a preform 242 for a second spar segment 114, the second spar segment 114 comprising a splice region 341, the splice regions 341 being arranged in series with one another in a spliced relationship; and a splice reinforcement 340 preform, the preforms 242 and the splice reinforcement 340 preform being simultaneously cured, the splice reinforcement 340 preform simultaneously covering at least a portion of the splice regions 341 to form a portion of the spar part 110.

[0131] Clause 32. The spar part 110 of Clause 31, wherein the splice reinforcement 340 is disposed at a rib intersection 360 of the spar part 110.

[0132] Clause 33. The spar part 110 according to Clause 31 or 32, wherein a splice reinforcement 340 is disposed at a side of the spar segment 112, 114 opposite a rib intersection point 360 of the spar part 110, the rib intersection point 360 and the splice reinforcement 340 sandwiching a splice area 341.

[0133] Clause 34. The spar part 110 according to Clause 31, 32, or 33, wherein the splice area 341 between the preforms 242 defines at least one of a lap splice, a butt splice, and a scarf splice.

[0134] Clause 35. The spar part 110 according to any of Clauses 31 to 34, wherein the preforms 242 and the splice reinforcement 340 preform include dry fibers infused with resin while covering at least a portion of the splice area 341.

[0135] Clause 36. The spar part 110 according to any of Clauses 31 to 35, further comprising: a preform 242 for a third spar segment 116, the third spar segment 116 including a splice area 341-1; and a second splice reinforcement 340-1 preform covering at least a portion of the splice area 341-1 of the third spar segment 116 and a portion of the second splice area 341-1 of the second spar segment 114.

[0136] Clause 37. Manufacturing a portion of an aircraft 10 using the spar part 110 of any of Clauses 31 to 36.

[0137] Clause 38. An aircraft wing 12, the wing comprising: a first spar segment 112 comprising a fiber-reinforced material 301, the first spar segment 112 comprising a splice region 341; a second spar segment 114 comprising a fiber-reinforced material 301-1, the second spar segment 114 comprising a first splice region 341 and a second splice region 341-1, the first splice region 341 of the second spar segment 114 being disposed in series with the splice region 341 of the first spar segment 112; a splice reinforcement 340 covering at least a portion of the splice region 341 of the first spar segment 112 and at least a portion of the first splice region 341 of the second spar segment 114, the splice reinforcement 340 being co-cured with the first spar segment 112 and the second spar segment 114; a third spar segment 116 comprising a fiber-reinforced material 301-2, the third spar segment 116 comprising a splice region 341-1, the splice region 341-1 being disposed in series with the second splice region 341-1 of the second spar segment 114; and a second splice reinforcement 340-1 covering at least a portion of the splice region 341-1 of the third spar segment 116 and at least a portion of the second splice region 341-1 of the second spar segment 114, the second splice reinforcement 340-1 being co-cured with the second spar segment 114 and the third spar segment 116.

[0138] Clause 39. The aircraft wing 12 of Clause 38, wherein at least one of the splice reinforcement 340 and the second splice reinforcement 340-1 extends across a rib intersection point 360 of the spar part 110.

[0139] Clause 40. The aircraft wing 12 of Clause 38, wherein prior to co-curing, the first spar segment 112, the second spar segment 114, the third spar segment 116, the splice reinforcement 340, and the second splice reinforcement 340-1 comprise dry fibers, the dry fibers being infused in place with a resin prior to co-curing.

[0140] Clause 41. A method 600 for manufacturing a spar part 110 for an aircraft 10, the method 600 comprising: manufacturing 602 a preform 242 of a fiber-reinforced material 301, 301-1, 301-2 for a spar segment 112, 114, 116; hardening 604 the preform 242 to form the spar segment 112, 114, 116; and applying 606 a fastener 278 that couples the spar segments 112, 114, 116 together to form the complete spar part 110.

[0141] Clause 42. The method 600 of any of Clauses 41, wherein applying 606 the fastener 278 includes applying a splice reinforcement 340 that partially covers the first spar segment 112 and the second spar segment 114; and installing the fastener 278 through the splice reinforcement 340 into the first spar segment 112 and the second spar segment 114.

[0142] Clause 43. The method 600 of Clause 42, wherein applying the splice reinforcement 340 includes clamping the first spar segment 112 and the second spar segment 114 between the splice reinforcement 340.

[0143] Clause 44. The method 600 of Clause 42, wherein applying the splice reinforcement 340 includes placing the splice reinforcement 340 at a location where the splice reinforcement 340 extends across a rib intersection 360 of the spar part 110.

[0144] Clause 45. The method 600 of Clause 42, wherein applying the splice reinforcement 340 includes placing the splice reinforcement 340 at a location where the splice reinforcement 340 extends across a side of the spar part 110 opposite the rib intersection 360 of the spar part 110.

[0145] Clause 46. The method 600 of Clause 42, further comprising: applying a second splice reinforcement 340-1 that partially covers the second spar segment 114 and a third spar segment 116; and installing the fastener 278 through the second splice reinforcement 340-1 into the second spar segment 114 and the third spar segment 116.

[0146] Clause 47. The method 600 of any of Clauses 41-46, wherein manufacturing 602 the preform 242 includes: dispensing a fiber-reinforced material 301 tow that forms a flat pack 232; and shaping the flat pack 232 into the preform 242 having a desired cross-sectional shape.

[0147] Clause 48. The method 600 of any of Clauses 41-47, wherein manufacturing 602 the preform 242 includes: dispensing a fiber-reinforced material 301 tow that forms a flat pack 232; combining a plurality of the flat packs 232 together into a full flat pack 233; and shaping the full flat pack 233 into the preform 242 having a desired cross-sectional shape.

[0148] Clause 49. A portion of an aircraft 10 assembled according to the method 600 of any of Clauses 41-48.

[0149] Clause 50. A spar part 110 for an aircraft 10, the spar part 110 comprising: a first spar segment 112 comprising a fiber reinforced material 301; a second spar segment 114 comprising a fiber reinforced material 301-1 and arranged in series with the first spar segment 112; a splice reinforcement 340 covering a splice area 341 between the first spar segment 112 and the second spar segment 114; and fasteners 278 installed through the splice reinforcement 340, the first spar segment 112, and the second spar segment 114 to form at least a portion of the spar part 110.

[0150] Clause 51. The spar part 110 of clause 50, wherein the splice reinforcement 340 is disposed on the first spar segment 112 and the second spar segment 114 across a rib intersection point 360.

[0151] Clause 52. The spar part 110 of clause 50 or 51, wherein the splice reinforcement 340 is disposed on an opposite side of the first spar segment 112 and the second spar segment 114 from the rib intersection point 360.

[0152] Clause 53. The spar part 110 of clause 50, 51, or 52, further comprising: a third spar segment 116 comprising a fiber reinforced material 301-2 and arranged in series with the second spar segment 114; a second splice reinforcement 340-1 covering a splice area 341-1 between the second spar segment 114 and the third spar segment 116; and fasteners 278 installed through the second splice reinforcement 340-1, the second spar segment 114, and the third spar segment 116 to form at least a portion of the spar part 110.

[0153] Clause 54. The spar part 110 of any of clauses 50-53, wherein the first spar segment 112 and the second spar segment 114 are sandwiched between the splice reinforcement 340.

[0154] Clause 55. Manufacturing a portion of an aircraft 10 using the spar part 110 of clause 50.

[0155] Clause 56. An aircraft wing 12, the wing comprising: a first spar segment 112 comprising a fiber reinforced material 301, the first spar segment 112 comprising a splice region 341; a second spar segment 114 comprising a fiber reinforced material 301, the second spar segment 114 comprising a first splice region 341 and a second splice region 341-1, the first splice region 341 of the second spar segment 114 being disposed in-line with the splice region 341 of the first spar segment 112; a splice reinforcement 340 covering at least a portion of the splice region 341 of the first spar segment 112 and at least a portion of the first splice region 341 of the second spar segment 114; and fasteners 278 installed through the splice reinforcement 340, the splice region 341 of the first spar segment 112, and the first splice region 341 of the second spar segment 114 to form at least a portion of a spar part 110.

[0156] Clause 57. The aircraft wing 12 of Clause 56, further comprising: a third spar segment 116 comprising a fiber reinforced material 301, the third spar segment 116 comprising a splice region 341-1 disposed in-line with the second splice region 341-1 of the second spar segment 114; a second splice reinforcement 340-1 covering at least a portion of the splice region 341-1 of the third spar segment 116 and at least a portion of the second splice region 341-1 of the second spar segment 114; and fasteners 278 installed through the second splice reinforcement 340-1, the second splice region 341-1 of the second spar segment 114, and the splice region 341-1 of the third spar segment 116 to form at least a portion of the spar part 110.

[0157] Clause 58. The aircraft wing 12 of Clause 57, wherein at least one of the splice reinforcement 340 and the second splice reinforcement 340-1 extends across a rib intersection 360.

[0158] Clause 59. The aircraft wing of either of Claims 57 or 58, wherein at least one of: the splice reinforcement 340 is disposed on a side of the first spar segment 112 and the second spar segment 114 opposite the rib intersection 360; and the second splice reinforcement 340-1 is disposed on a side of the second spar segment 114 and the third spar segment 116 opposite the rib intersection 360.

[0159] Clause 60. Manufacturing a portion of an aircraft 10 using the aircraft wing 12 of any of Clauses 57-59.

[0160] Clause 61. A method 700 for manufacturing a spar for an aircraft 10, the method comprising: manufacturing 702 preforms 242 of fiber-reinforced material 301 for spar segments 112, 114, 116, at least one of the spar segments 112, 114, 116 including a kink 370, each kink 370 being entirely contained within a preform 242; hardening 704 the preforms 242 to form the spar segments 112, 114, 116; and assembling 706 the spar segments 112, 114, 116 together to form a complete spar part 110 exhibiting at least one of the kinks 370.

[0161] Clause 62. The method 700 of clause 61, wherein each kink 370 includes an inflection point 381 at which there is an intersection of a first neutral axis 350 and a second neutral axis 350-1 of the spar.

[0162] Clause 63. The method 700 of clause 61 or 62, wherein manufacturing 702 the preforms 242 includes changing an axial direction of the preform 242 at each kink 370.

[0163] Clause 64. The method 700 of clause 61, 62, or 63, wherein manufacturing 702 the preforms 242 includes changing an axial direction of the preform 242 at each kink 370, the change in axial direction being a kink angle between 2 degrees and 10 degrees.

[0164] Clause 65. The method 700 of any of clauses 61-64, wherein each preform 242 includes a first end and a second end opposite the first end, and manufacturing 702 the preforms 242 includes placing each kink 370 at least about 30 cm (1 foot) from an end of the preform 242.

[0165] Clause 66. The method 700 of any of clauses 61-65, wherein each preform 242 includes at least one splice region 341, 341-1, and assembling 706 the spar segments 112, 114, 116 includes applying splice reinforcements 340, 340-1 to the spar segments 112, 114, 116 within the splice regions 341, 341-1.

[0166] Clause 67. The method 700 of clause 66, wherein applying the splice reinforcements 340, 340-1 to the spar segments 112, 114, 116 includes: forming at least one of a lap splice, a butt splice, and a scarf splice between adjacent splice regions 341, 341-1; and attaching the splice reinforcements 341, 341-1 to the splice regions 340, 340-1 using one of co-curing, co-bonding, and fasteners 278.

[0167] Clause 68. The method 700 of clause 66 or 67, wherein applying the splice reinforcement 340, 340-1 to the spar segment 112, 114, 116 within the splice region 341, 341-1 includes disposing the splice reinforcement 340, 340-1 across a rib intersection point 360 on the spar.

[0168] Clause 69. The method 700 of clause 66, 67, or 68, wherein applying the splice reinforcement 340, 340-1 to the spar segment 112, 114, 116 within the splice region 341, 341-1 includes sandwiching the splice region 341, 341-1 between the splice reinforcement 340, 340-1 and the rib.

[0169] Clause 70. The method 700 of any of clauses 66 to 69, wherein applying the splice reinforcement 340, 340-1 to the spar segment 112, 114, 116 within the splice region 341, 341-1 includes sandwiching the splice region 341, 341-1 between a forward splice reinforcement and an aft splice reinforcement.

[0170] Clause 71. The method 700 of any of clauses 61 to 70, wherein each preform 242 includes at least one splice region 341, 341-1, and manufacturing the preform 242 of fiber- reinforced material 301 for the spar segment 112, 114, 116 includes placing a kink 370 outside the splice region 341, 341-1 of the spar segment 112, 114, 116.

[0171] Clause 72. An aircraft spar part 110, comprising: a first spar segment 112 including a fiber-reinforced material 301, at least one splice region 341; a second spar segment 114 including a fiber-reinforced material 301, at least one splice region 341, and a kink 370 outside the splice region 341, wherein the respective splice regions 341 are disposed in series with each other; and a splice reinforcement 340 covering at least a portion of the splice region 341 of the first spar segment 112 and at least a portion of the corresponding splice region 341 of the second spar segment 114, the splice reinforcement 340 coupled to the first spar segment 112 and the second spar segment 114.

[0172] Clause 73. The aircraft spar part 110 of clause 72, wherein the kink 370 includes a change in an axial direction of the spar part 110.

[0173] Clause 74. The aircraft spar part 110 of Clause 73, wherein the change in axial direction is between 2 and 10 degrees.

[0174] Clause 75. The aircraft spar part 110 of Clause 72, 73, or 74, wherein the kink 370 is spaced more than about 30 cm (1 foot) from the end of the second spar segment 114.

[0175] Clause 76. The aircraft spar part 110 of any of Clauses 72 to 75, wherein the first spar segment 112 includes the kink 370 outside the splice region 341.

[0176] Clause 77. The aircraft spar part 110 of any of Clauses 72 to 76, wherein the splice stiffener 340 is positioned across the rib intersection 360.

[0177] Clause 78. The aircraft spar part 110 of Clause 77, wherein the splice stiffener 340 comprises one of: a front splice stiffener and a rear splice stiffener sandwiching the splice region 341; and a splice stiffener 340 placed to sandwich the splice region 341 between the splice stiffener 340 and the rib.

[0178] Clause 79. The aircraft spar part 110 of any of Clauses 72 to 78, further comprising: a third spar segment 116 including the fiber reinforced material 301, at least one splice region 341-1 and a kink 370 outside the splice region 341-1, the splice regions 341-1 being disposed in series with corresponding splice regions 341-1 of the second spar segment 114; and a second splice stiffener 340-1 covering at least a portion of the splice regions 341-1 of the third spar segment 116 and at least a portion of the corresponding splice regions 341-1 of the second spar segment 114, the second splice stiffener 340-1 being coupled with the second spar segment 114 and the third spar segment 116.

[0179] Clause 80. An aircraft wing 12, the aircraft wing comprising: a first spar segment 112 comprising a fiber-reinforced material 301, the first spar segment 112 comprising a splice region 341; a second spar segment 114 comprising a fiber-reinforced material 301, the second spar segment 114 comprising a first splice region 341 and a second splice region 341-1, the first splice region 341 of the second spar segment 114 being disposed in-line with the splice region 341 of the first spar segment 112; a splice reinforcement 340 covering at least a portion of the splice region 341 of the first spar segment 112 and at least a portion of the first splice region 341 of the second spar segment 114; a third spar segment 116 comprising a fiber-reinforced material 301, the third spar segment 116 comprising a splice region 341-1 disposed in-line with the second splice region 341-1 of the second spar segment 114; a second splice reinforcement 340-1 covering at least a portion of the splice region 341-1 of the third spar segment 116 and at least a portion of the second splice region 341-1 of the second spar segment 114; and at least one inflection 370 in one or more of the first spar segment 112, the second spar segment 114, and the third spar segment 116, the at least one inflection 370 being external to the splice regions 341, 341-1 and comprising a change in an axial direction of the spar segments 112, 114, 116 at a location of the inflection 370.

[0180] Clause 81. A method 800 for manufacturing a spar part 110 for an aircraft 10, the method 800 comprising: manufacturing 802 a preform 242 for a first spar segment 112, the preform 242 comprising a sub-inflection 371-1 proximate one end of the first spar segment 112; manufacturing 802 a preform 242 for a second spar segment 114, the preform 242 comprising a sub-inflection 371 proximate one end of the second spar segment 114; aligning 804 the ends of the preforms 242 such that the sub-inflections 371-1, 371 are proximate to one another within a splice region 341; and joining 806 the spar segments 112, 114 together in the splice region 341 to form at least a portion of the spar part 110 exhibiting an inflection 370.

[0181] Clause 82. The method 800 of clause 81, wherein joining 806 the spar segments 112, 114 comprises joining the spar segments 112, 114 in the splice region 341 using at least one of co-curing, co-bonding, installation of a splice reinforcement 340, and installation of a fastener 278.

[0182] Clause 83. The method 800 of any of clauses 81 or 82, wherein manufacturing 802 the preform 242 includes manufacturing 802 the preform 242 such that each sub-kink 371-1, 371 has an equal angular offset.

[0183] Clause 84. The method 800 of any of clauses 81, 82, or 83, wherein manufacturing 802 the preform 242 includes manufacturing 802 the preform 242 such that the sub-kinks 371-1, 371 have unequal angular offsets.

[0184] Clause 85. The method 800 of any of clauses 81 to 84, wherein manufacturing 802 the preform 242 includes manufacturing 802 the preform 242 such that the sub-kinks 371-1, 371 together change the axial direction of the spar part 110 by between 2 degrees and 10 degrees.

[0185] Clause 86. The method 800 of any of clauses 81 to 85, wherein aligning 804 the ends of the preform 242 includes forming one of a lap joint, a butt joint, and a scarf joint with the ends of the preform 242 in the splice region 341.

[0186] Clause 87. The method 800 of any of clauses 81 to 86, wherein joining 806 the spar segments 112, 114 together includes applying a splice reinforcement 340 to the splice region 341 opposite the rib intersection point 360.

[0187] Clause 88. The method 800 of any of clauses 81 to 87, wherein joining 806 the spar segments 112, 114 together includes sandwiching the sub-kinks 371-1, 371 between the splice reinforcement 340 and the rib.

[0188] Clause 89. The method 800 of any of clauses 81 to 88, wherein joining 806 the spar segments 112, 114 together includes sandwiching the sub-kinks 371-1, 371 between a front splice reinforcement and a back splice reinforcement.

[0189] Clause 90. The method 800 of any of clauses 81 to 89, wherein the method 800 assembles a portion of the aircraft 10.

[0190] Clause 91. A spar part 110 for an aircraft 10, the spar part 110 comprising: a first spar segment 112 comprising a fiber-reinforced material 301 and including a sub-kink 371-1 disposed at an end; a second spar segment 114 comprising a fiber-reinforced material 301 and including a sub-kink 371 disposed at an end, the end of the first spar segment 112 having the sub-kink 371-1 adjacent to the end of the second spar segment 114 having the sub-kink 371 such that the sub-kinks 371-1, 371 together form a kink 370, and the ends define a splice region 341; and a splice reinforcement 340 structurally integrating the first and second spar segments 112, 114 within the splice region 341.

[0191] Clause 92. The spar part 110 of clause 91, wherein the sub-kink 371-1 associated with the first spar segment 112 is disposed within about 30 cm (1 foot) of the end of the first spar segment 112; and the sub-kink 371 associated with the second spar segment 114 is disposed within about 30 cm (1 foot) of the end of the second spar segment 114.

[0192] Clause 93. The spar part 110 of clause 91 or 92, wherein each sub-kink 371-1, 371 changes an axial direction of the spar part 110 by half the amount of the kink 370.

[0193] Clause 94. The spar part 110 of clause 91, 92, or 93, wherein the sub-kinks 371-1, 371 together change an axial direction of the spar part 110 between 2 and 10 degrees.

[0194] Clause 95. The spar part 110 of any of clauses 91-94, wherein the splice reinforcement 340 is disposed opposite a rib intersection point 360 defined for the spar part 110.

[0195] Clause 96. The spar part 110 of any of clauses 91-95, wherein the splice reinforcement 340 is arranged to sandwich the splice region 341 and the sub-kinks 371-1, 371.

[0196] Clause 97. The spar part 110 of any of clauses 91-96, wherein to structurally integrate the first and second spar segments 112, 114, the splice reinforcement 340 is one of: co-bonded to, co-cured to, and attached to the first and second spar segments 112, 114 with fasteners.

[0197] Clause 98. An aircraft 10 manufactured using the spar part 110 of any of Clauses 91-97.

[0198] Clause 99. An aircraft wing 12 comprising: a first spar segment 112 comprising a fiber reinforced material 301 and comprising a sub-kink 371-1 disposed at an end; a second spar segment 114 comprising a fiber reinforced material 301 and comprising a sub-kink 371 disposed at an end, the end of the first spar segment 112 having the sub-kink 371-1 adjacent to the end of the second spar segment 114 having the sub-kink 371 such that the sub-kinks 371-1, 371 together form a kink 370, and the ends define a splice region 341; a splice reinforcement 340 structurally integrating the first spar segment 112 and the second spar segment 114 within the splice region 341; a third spar segment 116 comprising a fiber reinforced material 301, the third spar segment 116 comprising a splice region 341-1 disposed in series with a second splice region 341-1 at an opposite end of the second spar segment 114; and a second splice reinforcement 340-1 covering at least a portion of the splice region 341-1 of the third spar segment 116 and at least a portion of the second splice region 341-1 of the second spar segment 114.

[0199] Clause 100. The aircraft wing 12 of Clause 99, wherein the splice reinforcement 340 is one of: co-bonded to, co-cured to, and attached to the first spar segment 112 and the second spar segment 114 with fasteners.

[0200] Any of the various control elements (e.g., electrical or electronic components) shown in the drawings or described herein can be implemented as hardware, software implemented processor, software implemented processor, or some combination of these. For example, an element can be implemented as dedicated hardware. A dedicated hardware element can be referred to as a "processor", "controller", or some similar terminology. When provided by a processor, the functions can be provided by a single dedicated processor, by a single shared processor, or by a plurality of individual processors, some of which can be shared. Moreover, explicit use of the term "processor" or "controller" should not be construed to refer exclusively to hardware capable of executing software, and can implicitly include, without limitation, digital signal processor (DSP) hardware, network processor, application specific integrated circuit (ASIC), or other circuitry, field programmable gate array (FPGA), read only memory (ROM) for storing software, random access memory (RAM), nonvolatile storage medium, logic or some other physical hardware component or module.

[0201] Further, a control element can be implemented as instructions executable by a processor or computer to perform the function of the element. Some examples of instructions are software, program code, and firmware. The instructions are operable when executed by the processor to direct the processor to perform the function of the element. The instructions can be stored on a storage device readable by the processor. Some examples of storage devices are digital or solid state memory, magnetic storage media such as disks and tapes, hard drives, or optically readable digital data storage media.

[0202] Although specific embodiments are described herein, the scope of the disclosure is not limited to those specific embodiments. The scope of the disclosure is defined by the following claims.

[0203] The present disclosure also includes the following clauses.

[0204] Clause

[0205] 1. A method for manufacturing a spar part for an aircraft, the method comprising:

[0206] manufacturing a preform of fiber reinforced material for a spar segment;

[0207] hardening the preform to form the spar segment; and

[0208] joining the spar segments together to form the complete spar part.

[0209] 2. The method of clause 1, wherein joining the spar segments together comprises:

[0210] applying a splice reinforcement that partially covers the first spar segment and the second spar segment; and

[0211] joining the splice reinforcement to the first spar segment and the second spar segment.

[0212] 3. The method of clause 2, wherein applying the splice reinforcement comprises clamping the splice reinforcement between a splice area of the first spar segment and a splice area of the second spar segment.

[0213] 4. The method of clause 2, wherein applying the splice reinforcement comprises clamping the splice area of the first spar segment and the splice area of the second spar segment between a front splice reinforcement and a back splice reinforcement.

[0214] 5. The method of clause 2, wherein applying the splice reinforcement comprises placing the splice reinforcement at a location where the splice reinforcement extends across a rib intersection point of the spar part.

[0215] 6. The method of clause 2, wherein applying the splice reinforcement comprises clamping the first spar segment and the second spar segment between the splice reinforcement and the rib.

[0216] 7. The method of clause 2, wherein bonding the splice reinforcement comprises hardening the splice reinforcement by co-bonding in an extrusion kettle.

[0217] 8. The method of clause 1, wherein bonding the spar segments together comprises:

[0218] applying one of an adhesive and a resin to the spar segments; and

[0219] forming a splice between the spar segments in the splice region of the spar segments.

[0220] 9. The method of clause 1, wherein bonding the spar segments together comprises forming one of a butt splice, a lap splice, and a scarf splice between the splice region of the first spar segment and the splice region of the second spar segment.

[0221] 10. A spar for an aircraft, the spar comprising:

[0222] a first spar segment comprising a fiber reinforced material, the first spar segment comprising a splice region;

[0223] a second spar segment comprising a fiber reinforced material, the second spar segment comprising a first splice region and being disposed in series with the first spar segment; and

[0224] a splice reinforcement covering at least a portion of the first splice region of the second spar segment and the splice region of the first spar segment, the splice reinforcement being bonded to the first spar segment and the second spar segment.

[0225] 11. The spar of clause 10, wherein the spar is one of a front spar 110-1 and a rear spar 110-2.

[0226] 12. The spar of clause 10, wherein the splice reinforcement is clamped between the splice region of the first spar segment and the first splice region of the second spar segment.

[0227] 13. The spar of clause 10, wherein the splice reinforcement comprises a front splice reinforcement and a rear splice reinforcement, the splice region of the first spar segment and the first splice region of the second spar segment being clamped between the front splice reinforcement and the rear splice reinforcement.

[0228] 14. The spar of clause 10, wherein the splice reinforcement extends across a rib intersection of the spar.

[0229] 15. The spar of clause 10, wherein the splice region of the first spar segment and the first splice region of the second spar segment are sandwiched between the splice stiffener and the rib.

[0230] 16. The spar of clause 10, wherein the splice stiffener is co-bonded to the splice region of the first spar segment and the first splice region of the second spar segment.

[0231] 17. The spar of clause 10, wherein the second spar segment further comprises a second splice region, the spar further comprising:

[0232] a third spar segment comprising a fiber-reinforced material, the third spar segment comprising a splice region and being disposed in series with the second spar segment; and

[0233] a second splice stiffener covering at least a portion of the splice region of the third spar segment and at least a portion of the second splice region of the second spar segment, the second splice stiffener being co-bonded to the second spar segment and the third spar segment.

[0234] 18. Using the spar of clause 10 to manufacture a portion of an aircraft.

[0235] 19. An aircraft wing comprising a multi-segment spar part, the spar part comprising:

[0236] a first spar segment comprising a fiber-reinforced material, the first spar segment comprising a splice region;

[0237] a second spar segment comprising a fiber-reinforced material, the second spar segment comprising a first splice region, the second spar segment being disposed in series with the first spar segment;

[0238] a splice stiffener covering at least a portion of the first splice region of the second spar segment and the splice region of the first spar segment, the splice stiffener being co-bonded to the first spar segment and the second spar segment;

[0239] a third spar segment comprising a fiber-reinforced material, the third spar segment comprising a splice region and being disposed in series with the second spar segment; and

[0240] a second splice stiffener covering at least a portion of the splice region of the third spar segment and at least a portion of a second splice region of the second spar segment, the second splice stiffener being co-bonded to the second spar segment and the third spar segment.

[0241] 20. The aircraft wing of clause 19, wherein at least one of the splice stiffener and the second splice stiffener extends across a rib intersection of the multi-segment spar part.

[0242] co-curing

[0243] 21. A method for manufacturing a spar part for an aircraft, the method comprising:

[0244] manufacturing preforms of fiber-reinforced material for first and second spar segments;

[0245] splicing an end of the preform of the first spar segment to an end of the preform of the second spar segment to define a splice region;

[0246] applying at least one preform for a splice reinforcement to the splice region; and

[0247] simultaneously hardening the preforms for the spar segments and the splice reinforcement to form a portion of the spar part.

[0248] 22. The method of clause 21, wherein simultaneously hardening the preforms for the spar segments and the preform for the splice reinforcement comprises:

[0249] vacuum bagging the preforms for the spar segments and the preform for the splice reinforcement with a vacuum bag;

[0250] consolidating the preforms for the spar segments and the preform for the splice reinforcement via the vacuum bag; and

[0251] heating the preforms for the spar segments and the preform for the splice reinforcement.

[0252] 23. The method of clause 21, wherein:

[0253] applying at least one preform for a splice reinforcement comprises picking up and laying down the preform for the splice reinforcement such that the splice reinforcement extends across a rib intersection of the spar part.

[0254] 24. The method of clause 21, wherein:

[0255] applying at least one preform for a splice reinforcement comprises laying the preform for the splice reinforcement onto the spliced ends of the first and second spar segments such that the splice reinforcement extends across a rib intersection of the spar part.

[0256] 25. The method of clause 21, wherein:

[0257] applying at least one preform for a splice reinforcement comprises laying the preform for the splice reinforcement across a side of the first and second spar segments opposite a rib intersection of the spar part.

[0258] 26. The method of clause 21, wherein simultaneously hardening the preforms for the spar segment and the splice reinforcement comprises:

[0259] arranging dry fiber preforms for the first spar segment, the second spar segment, and the splice reinforcement; and

[0260] infusing the dry fibers with resin.

[0261] 27. The method of clause 21, wherein splicing the end of the first spar segment preform to the end of the second spar segment preform to define a splice region comprises splicing the ends with at least one of a lap splice, a butt splice, and a scarf splice.

[0262] 28. The method of clause 21, wherein manufacturing the preforms comprises:

[0263] dispensing fiber reinforcement tows to form a flat pack; and

[0264] shaping the flat pack into a preform having a desired cross-sectional shape.

[0265] 29. The method of clause 21, wherein manufacturing the preforms comprises:

[0266] dispensing fiber reinforcement tows to form a flat pack;

[0267] combining a plurality of flat packs together into a full flat pack; and

[0268] shaping the full flat pack into a preform having a desired cross-sectional shape.

[0269] 30. A portion of an aircraft assembled according to the method of clause 21.

[0270] 31. A spar part, the spar part comprising:

[0271] a preform for a first spar segment, the first spar segment comprising a splice region;

[0272] a preform for a second spar segment, the second spar segment comprising a splice region, the splice regions arranged in series with one another in a spliced relationship; and

[0273] a splice reinforcement preform, the preforms and the splice reinforcement preform being simultaneously cured, the splice reinforcement preform simultaneously covering at least a portion of the splice regions to form a portion of the spar part.

[0274] 32. The spar part of clause 31, wherein the splice reinforcement is disposed at a rib intersection of the spar part.

[0275] 33. The spar part of clause 31, wherein a splice reinforcement is disposed at an opposite side of the spar segment from a rib intersection point of the spar part, the rib intersection point and the splice reinforcement sandwiching the splice area.

[0276] 34. The spar part of clause 31, wherein the splice area between the preforms defines at least one of a lap splice, a butt splice, and a scarf splice.

[0277] 35. The spar part of clause 31, wherein the preforms and the splice reinforcement preform comprise dry fibers infused with resin while the splice reinforcement preform covers at least a portion of the splice area (341).

[0278] 36. The spar part of clause 31, further comprising:

[0279] a preform 242 for a third spar segment 330, the third spar segment comprising a splice area; and

[0280] a second splice reinforcement preform 242 covering at least a portion of the splice area of the third spar segment and a portion of the second splice area of the second spar segment.

[0281] 37. A method of manufacturing a portion of an aircraft using the spar part of clause 31.

[0282] 38. An aircraft wing, comprising:

[0283] a first spar segment comprising a fiber reinforced material, the first spar segment comprising a splice area;

[0284] a second spar segment comprising a fiber reinforced material, the second spar segment comprising a first splice area and a second splice area, the first splice area of the second spar segment disposed in series with the splice area of the first spar segment;

[0285] a splice reinforcement covering at least a portion of the splice area of the first spar segment and at least a portion of the first splice area of the second spar segment, the splice reinforcement co-cured with the first spar segment and the second spar segment;

[0286] a third spar segment comprising a fiber reinforced material, the third spar segment comprising a splice area disposed in series with the second splice area of the second spar segment; and

[0287] a second splice reinforcement covering at least a portion of the splice region of the third spar segment and at least a portion of the second splice region of the second spar segment, the second splice reinforcement being co-cured with the second spar segment and the third spar segment.

[0288] 39. The aircraft wing of clause 38, wherein at least one of the splice reinforcement and the second splice reinforcement extends across a rib intersection of the spar part.

[0289] 40. The aircraft wing of clause 38, wherein prior to co-curing, the first spar segment, the second spar segment, the third spar segment, the splice reinforcement, and the second splice reinforcement comprise dry fibers that are infused in place with resin prior to co-curing.

[0290] fastener

[0291] 41. A method for manufacturing a spar part for an aircraft, the method comprising:

[0292] manufacturing a preform of fiber-reinforced material for a spar segment;

[0293] hardening the preform to form the spar segment; and

[0294] applying fasteners that couple the spar segments together to form the complete spar part.

[0295] 42. The method of clause 41, wherein applying the fasteners comprises:

[0296] applying a splice reinforcement that partially covers the first spar segment and the second spar segment; and

[0297] installing the fasteners through the splice reinforcement into the first spar segment and the second spar segment.

[0298] 43. The method of clause 42, wherein applying the splice reinforcement comprises clamping the first spar segment and the second spar segment between the splice reinforcement.

[0299] 44. The method of clause 42, wherein applying the splice reinforcement comprises placing the splice reinforcement at a location where the splice reinforcement extends across a rib intersection of the spar part.

[0300] 45. The method of clause 42, wherein applying the splice reinforcement comprises placing the splice reinforcement at a location where the splice reinforcement extends across a side of the spar part opposite a rib intersection of the spar part.

[0301] 46. The method of clause 42, further comprising:

[0302] a second splice reinforcement that partially covers the second spar segment and the third spar segment; and

[0303] installing fasteners through the second splice reinforcement into the second spar segment and the third spar segment.

[0304] 47. The method of clause 41, wherein manufacturing the preform comprises:

[0305] dispensing a fiber-reinforced material tows that form a flat pack; and

[0306] shaping the flat pack into a preform having a desired cross-sectional shape.

[0307] 48. The method of clause 41, wherein manufacturing the preform comprises:

[0308] dispensing a fiber-reinforced material tows that form a flat pack;

[0309] combining the plurality of flat packs together into a full flat pack; and

[0310] shaping the full flat pack into a preform having a desired cross-sectional shape.

[0311] 49. A portion of an aircraft assembled according to the method of any one of clauses 41 to 48.

[0312] 50. A spar part for an aircraft, the spar part comprising:

[0313] a first spar segment comprising a fiber-reinforced material;

[0314] a second spar segment comprising a fiber-reinforced material and arranged in series with the first spar segment;

[0315] a splice reinforcement covering a splice area between the first spar segment and the second spar segment; and

[0316] fasteners installed through the splice reinforcement, the first spar segment, and the second spar segment to form at least a portion of the spar part.

[0317] 51. The spar part of clause 50, wherein the splice reinforcement is disposed on the first spar segment and the second spar segment across a rib intersection point.

[0318] 52. The spar part of clause 50, wherein the splice reinforcement is disposed on a side of the first spar segment and the second spar segment opposite the rib intersection point.

[0319] 53. The spar part of clause 50, further comprising:

[0320] a third spar segment comprising fiber-reinforced material and disposed in series with the second spar segment;

[0321] a second splice reinforcement covering a splice area between the second spar segment and the third spar segment; and

[0322] fasteners installed through the second splice reinforcement, the second spar segment, and the third spar segment to form at least a portion of a spar part.

[0323] 54. The spar part of clause 50, wherein the first spar segment and the second spar segment are sandwiched between the splice reinforcements.

[0324] 55. A method of manufacturing a portion of an aircraft using the spar part of any one of clauses 50 to 54.

[0325] 56. An aircraft wing, the wing comprising:

[0326] a first spar segment comprising fiber-reinforced material, the first spar segment comprising a splice area;

[0327] a second spar segment comprising fiber-reinforced material, the second spar segment comprising a first splice area and a second splice area, the first splice area of the second spar segment disposed in series with the splice area of the first spar segment;

[0328] a splice reinforcement covering at least a portion of the splice area of the first spar segment and at least a portion of the first splice area of the second spar segment; and

[0329] fasteners installed through the splice reinforcement, the splice area of the first spar segment, and the first splice area of the second spar segment to form at least a portion of a spar part.

[0330] 57. The aircraft wing of clause 56, the wing further comprising:

[0331] a third spar segment comprising fiber-reinforced material, the third spar segment comprising a splice area disposed in series with the second splice area of the second spar segment;

[0332] a second splice reinforcement covering at least a portion of the splice area of the third spar segment and at least a portion of the second splice area of the second spar segment; and

[0333] fasteners installed through the second splice reinforcement, the second splice area of the second spar segment, and the splice area of the third spar segment to form at least a portion of a spar part.

[0334] 58. The aircraft wing of Clause 57, wherein at least one of the splice stiffener and the second splice stiffener extends across the rib intersection.

[0335] 59. The aircraft wing of Claim 57, wherein at least one of:

[0336] the splice stiffener is disposed on a side of the first spar segment and the second spar segment opposite the rib intersection; and

[0337] the second splice stiffener is disposed on a side of the second spar segment and the third spar segment opposite the rib intersection.

[0338] 60. Manufacturing a portion of an aircraft using the aircraft wing of any one of Clauses 57 to 59.

[0339] spar kink

[0340] 61. A method for manufacturing a spar for an aircraft, the method comprising:

[0341] manufacturing preforms of fiber-reinforced material for spar segments, at least one of the spar segments including a kink, each kink being entirely contained within a preform;

[0342] hardening the preforms to form the spar segments; and

[0343] assembling the spar segments together to form a complete spar part exhibiting at least one of the kinks.

[0344] 62. The method of Clause 61, wherein each kink includes a cusp at which there is an intersection of a first neutral axis and a second neutral axis of the spar.

[0345] 63. The method of Clause 61, wherein manufacturing the preforms includes changing an axial direction of the preform at each kink.

[0346] 64. The method of Clause 61, wherein manufacturing the preforms includes changing an axial direction of the preform at each kink, the change in axial direction being a kink angle between 2 degrees and 10 degrees.

[0347] 65. The method of Clause 61, wherein each preform includes a first end and a second end opposite the first end, and manufacturing the preforms includes placing each kink at a location at least about 30 cm (1 foot) from an end of the preform.

[0348] 66. The method of clause 61, wherein each preform includes at least one splice region, and assembling the spar segments includes applying a splice reinforcement to the spar segments within the splice regions.

[0349] 67. The method of clause 66, wherein applying the splice reinforcement to the spar segments includes:

[0350] forming at least one of a lap splice, a butt splice, and a scarf splice between adjacent splice regions; and

[0351] attaching the splice reinforcement to the splice regions using one of co-curing, co-bonding, and fasteners.

[0352] 68. The method of clause 66, wherein applying the splice reinforcement to the spar segments within the splice regions includes arranging the splice reinforcement across rib on-rib intersections.

[0353] 69. The method of clause 66, wherein applying the splice reinforcement to the spar segments within the splice regions includes sandwiching the splice regions between the splice reinforcement and the ribs.

[0354] 70. The method of clause 66, wherein applying the splice reinforcement to the spar segments within the splice regions includes sandwiching the splice regions between a forward splice reinforcement and an aft splice reinforcement.

[0355] 71. The method of clause 61, wherein each preform includes at least one splice region, and manufacturing the preforms of fiber-reinforced material for the spar segments includes placing a kink outside the splice regions of the spar segments.

[0356] 72. An aircraft spar part, comprising:

[0357] a first spar segment including fiber-reinforced material and at least one splice region;

[0358] a second spar segment including fiber-reinforced material, at least one splice region, and a kink outside the splice region, wherein the respective splice regions are disposed in series with one another; and

[0359] a splice reinforcement covering at least a portion of the splice region of the first spar segment and at least a portion of the corresponding splice region of the second spar segment, the splice reinforcement coupled to the first spar segment and the second spar segment.

[0360] 73. The aircraft spar part of clause 72, wherein the kink includes a change in an axial direction of the spar part.

[0361] 74. The aircraft spar part of Clause 73, wherein the change in axial direction is between 2 and 10 degrees.

[0362] 75. The aircraft spar part of Clause 72, wherein the kink is spaced more than about 30 cm (1 foot) from the end of the second spar segment.

[0363] 76. The aircraft spar part of Clause 72, wherein the first spar segment includes the kink outside the splice region.

[0364] 77. The aircraft spar part of Clause 72, wherein the splice stiffener is positioned across a rib intersection.

[0365] 78. The aircraft spar part of Clause 77, wherein the splice stiffener comprises one of:

[0366] a front splice stiffener and a back splice stiffener sandwiching the splice region; and

[0367] a splice stiffener placed to sandwich the splice region between the splice stiffener and the rib.

[0368] 79. The aircraft spar part of Clause 72, further comprising:

[0369] a third spar segment comprising a fiber-reinforced material, at least one splice region, and a kink outside the splice region, the splice regions being arranged in series with the corresponding splice regions of the second spar segment; and

[0370] a second splice stiffener covering at least a portion of the splice regions of the third spar segment and at least a portion of the corresponding splice regions of the second spar segment, the second splice stiffener being coupled with the second spar segment and the third spar segment.

[0371] 80. An aircraft wing, comprising:

[0372] a first spar segment comprising a fiber-reinforced material, the first spar segment including a splice region;

[0373] a second spar segment comprising a fiber-reinforced material, the second spar segment including a first splice region and a second splice region, the first splice region of the second spar segment being arranged in series with the splice region of the first spar segment;

[0374] a splice stiffener covering at least a portion of the splice region of the first spar segment and at least a portion of the first splice region of the second spar segment;

[0375] a third spar segment comprising a fiber-reinforced material, the third spar segment comprising a splice region disposed in series with the second splice region of the second spar segment;

[0376] a second splice reinforcement covering at least a portion of the splice region of the third spar segment and at least a portion of the second splice region of the second spar segment; and

[0377] at least one inflection located in one or more of the first spar segment, the second spar segment, and the third spar segment, the at least one inflection being outside of the splice regions and comprising a change in an axial direction of the spar segment at a location of the inflection.

[0378] a separation ramp

[0379] 81. A method for manufacturing a spar part for an aircraft, the method comprising:

[0380] manufacturing a preform for a first spar segment, the preform comprising a sub- inflection proximate to one end of the first spar segment;

[0381] manufacturing a preform for a second spar segment, the preform comprising a sub- inflection proximate to one end of the second spar segment;

[0382] aligning the ends of the preforms such that the sub-inflections are proximate to each other within a splice region; and

[0383] joining the spar segments together in the splice region to form at least a portion of a spar part exhibiting an inflection.

[0384] 82. The method of clause 81, wherein joining the spar segments comprises joining the spar segments in the splice region using at least one of co-curing, co-bonding, installation of a splice reinforcement, and installation of a fastener.

[0385] 83. The method of clause 81, wherein manufacturing the preforms comprises manufacturing the preforms such that each sub-inflection has an equal angular offset.

[0386] 84. The method of clause 81, wherein manufacturing the preforms comprises manufacturing the preforms such that the sub-inflections have unequal angular offsets.

[0387] 85. The method of clause 81, wherein manufacturing the preforms comprises manufacturing the preforms such that the sub-inflections together change an axial direction of the spar part by between 2 degrees and 10 degrees.

[0388] 86. The method according to Clause 81, wherein aligning the ends of the preforms includes forming one of an lap splice, a butt splice, and an interlocking splice in the splicing area using the ends of the preforms.

[0389] 87. The method according to Clause 81, wherein joining these spar segments together includes applying splice reinforcement to the splice area opposite the rib intersection point.

[0390] 88. The method described in Clause 81, wherein joining these spar segments together includes clamping these sub-knocks between splice reinforcements and ribs.

[0391] 89. The method according to Clause 81, wherein joining these spar segments together includes clamping these sub-knocks between the front splice reinforcement and the rear splice reinforcement.

[0392] 90. A part of an aircraft assembled according to the method described in 81.

[0393] 91. A wing sparsity component for an aircraft, the wing sparsity component comprising:

[0394] The first wing segment includes fiber-reinforced material and includes a sub-knock at the end;

[0395] A second wing segment, comprising fiber-reinforced material and including sub-knocks at its ends, wherein the ends of the first wing segment having sub-knocks are adjacent to the ends of the second wing segment having sub-knocks, such that these sub-knocks together form a bend, and these ends define a splicing area; and

[0396] The splicing reinforcement integrates the first wing beam segment and the second wing beam segment structurally within the splicing area.

[0397] 92. The spar component as described in Clause 91, wherein:

[0398] The sub-bend associated with the first wing segment is located within approximately 30 cm (1 ft) of the end of the first wing segment; and

[0399] The sub-bend associated with the second wing segment is located within approximately 30 cm (1 ft) of the end of the second wing segment.

[0400] 93. The spar component as described in Clause 91, wherein each sub-bend changes the axial orientation of the spar component by half the amount of the bend.

[0401] 94. The spar component as described in Clause 91, wherein these sub-bends together change the axial orientation of the spar component between 2 degrees and 10 degrees.

[0402] 95. The spar component as described in Clause 91, wherein the splicing reinforcement is configured to be opposite to the intersection point of the rib defined for the spar component.

[0403] 96. The spar component as described in Clause 91, wherein the splicing reinforcement is arranged to sandwich the splicing area and the sub-knock in the middle.

[0404] 97. The spar component according to Clause 91, wherein, in order to structurally integrate the first spar segment and the second spar segment, the splicing reinforcement is one of the following: co-bonded to, co-cured to, and attached to the first spar segment and the second spar segment using fasteners.

[0405] 98. Use the spar parts described in Clause 91 to manufacture a part of an aircraft.

[0406] 99. An aircraft wing, the aircraft wing comprising:

[0407] The first wing segment includes fiber-reinforced material and includes a sub-knock at the end;

[0408] The second wing segment includes fiber-reinforced material and includes sub-bends at its ends, the ends of the first wing segment having sub-bends being adjacent to the ends of the second wing segment having sub-bends, such that the sub-bends together form a bend, and the ends defining a splicing area.

[0409] The splicing reinforcement integrates the first wing beam segment and the second wing beam segment structurally within the splicing area;

[0410] A third wing segment, comprising fiber-reinforced material, includes a splicing region connected in series with a second splicing region at the opposite end of a second wing segment; and

[0411] The second splicing reinforcement covers at least a portion of the splicing area of ​​the third wing segment and at least a portion of the second splicing area of ​​the second wing segment.

[0412] 100. The aircraft wing as described in Clause 99, wherein the splicing reinforcement is one of the following: co-bonded to, co-cured to, and attached to the first spar segment and the second spar segment using fasteners.

Claims

1. A method (700) for manufacturing a spar component for an aircraft (10), the method comprising: (702) A preform (242) of fiber-reinforced material (301) for spar segments (112, 114, 116) is manufactured, at least one of the spar segments (112, 114, 116) including a bend (370), each bend (370) being completely contained within the preform (242) and including a bend point (381), at which the bend angle between a first neutral axis (350) and a second neutral axis (350-1) of the preform (242) is between 2 degrees and 10 degrees; Harden (704) the preform (242) to form the spar segments (112, 114, 116); and The spar segments (112, 114, 116) are assembled (706) together to form a complete spar part (110) exhibiting at least one of the bends (370).

2. The method according to claim 1, wherein, The step of manufacturing (702) the preform includes changing the bend angle (370) of the preform (242) at each bend.

3. The method according to claim 1 or 2, wherein, Each preform (242) includes a first end and a second end opposite to the first end, and manufacturing (702) the preform (242) includes placing each bend (370) at a position at least 30 cm, or 1 foot, from the end of the preform.

4. The method according to claim 1, wherein, Each preform (242) includes at least one splicing area (341, 341-1), and assembling (706) the spar segments (112, 114, 116) includes applying splicing reinforcements (340, 340-1) to the spar segments (112, 114, 116) within the splicing areas (341, 341-1).

5. The method according to claim 4, wherein, The at least one splicing region (341, 341-1) exhibits a neutral axis aligned with the first neutral axis (350) or the second neutral axis (350-1).

6. The method according to claim 4 or 5, wherein, Applying the splice reinforcements (340, 340-1) to the spar segments (112, 114, 116) includes: At least one of lap splicing, butt splicing, and interlocking splicing is formed between adjacent ends of the wing beam segments (112, 114, 116); and The splicing reinforcement (340, 340-1) is attached to the splicing area (341, 341-1) using one of co-curing, co-bonding and fasteners.

7. The method according to claim 4 or 5, wherein, Applying splicing reinforcements (340, 340-1) to the spar segments (112, 114, 116) within the splicing areas (341, 341-1) includes one of the following: One of the splicing reinforcements (340, 340-1) is arranged across the rib intersection point (360) on the wing beam part; Each splicing area (341, 341-1) is sandwiched between the splicing reinforcement (340, 340-1) and the rib; The splicing area (341, 341-1) is sandwiched between the front splicing reinforcement and the rear splicing reinforcement.

8. The method according to claim 1 or 2, wherein, Each preform includes at least one splice area (341, 341-1), and the fabrication (702) of a preform (242) of fiber-reinforced material (301) for the spar segments (112, 114, 116) includes placing the bend (370) outside the splice area (341, 341-1) of the spar segments (112, 114, 116).

9. A wing spars component (110) for an aircraft, said wing spars component comprising: Includes fiber-reinforced material (301) and a first wing segment (112) at the outer end; It includes fiber-reinforced material (301), an inner end and a bend in the second wing segment (114), the bend including a bend point (381), at which the bend angle between the first neutral axis (350) and the second neutral axis (350-1) of the second wing segment (114) is between 2 degrees and 10 degrees; Wherein, the outer end of the first wing beam segment (112) and the inner end of the second wing beam segment (114) are connected in series to form a splicing area (341); and A splicing reinforcement (340) covers at least a portion of the first wing segment (112) and at least a portion of the second wing segment (114) in the splicing area (341), and the splicing reinforcement (340) is connected to the first wing segment (112) and the second wing segment (114).

10. The spar component according to claim 9, wherein, The splicing area (341) exhibits a neutral axis aligned with either the first neutral axis (350) or the second neutral axis (350-1).

11. The spar component according to claim 9 or 10, wherein, The bend (370) is separated from the inner end of the second wing segment (114) by more than 30 cm, or 1 foot.

12. The spar component according to claim 9 or 10, wherein, The first wing segment (112) includes a bend (370) located outside the splicing area (341).

13. The spar component according to claim 9 or 10, wherein, The splicing reinforcement (340) is positioned across the intersection point (360) of the ribs.

14. The spar component according to claim 9 or 10, further comprising: A third wing segment (116) comprising fiber-reinforced material (301), an inner end, and a bend (370), wherein the inner end of the third wing segment (116) is connected in series with the corresponding outer end of the second wing segment (114) to form a second splicing region (341-1), wherein the bend (370) is located outside the second splicing region (341-1); and The second splicing reinforcement (340-1) covers at least a portion of the third wing segment (116) and at least a portion of the second wing segment (114) in the second splicing area (341-1), and the second splicing reinforcement (340-1) is connected to the second wing segment (114) and the third wing segment (116).

15. The spar component according to claim 13, wherein, The splicing reinforcement (340) includes one of the following: The front splicing reinforcement and the rear splicing reinforcement sandwich the splicing area (341) between them; and The splicing reinforcement (340) is positioned to sandwich the splicing area (341) between the splicing reinforcement (340) and the rib.

16. An aircraft wing comprising at least one spars member according to any one of claims 9 to 15.

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