Manufacturing of multi-segment spurs

Abstract

To provide systems and methods for fabricating multiple-segment spars for an aircraft.SOLUTION: In one example, the method includes fabricating preforms of fiber-reinforced material for spar segments, hardening the preforms to form the spar segments, and bonding the spar segments together to form a completed spar detail. In addition to bonding, other examples include co-curing and fastening the spar segments. In additional examples, the spar segments include kinks or sub-kinks as described.SELECTED DRAWING: Figure 4

Description

Technical Field

[0001] The present disclosure relates to the field of manufacturing, and particularly to the manufacturing of composite parts of aircraft.

Background Art

[0002] The spar of an aircraft runs from the inboard direction to the outboard direction within the wing and provides structural strength to the wing. The spar of an aircraft can be manufactured from, for example, composite materials. Since the wing is a particularly long component of the aircraft, multiple spars extend the length of the wing. Due to the large size of the spar, the complexity and cost of dedicated equipment for layup, consolidation, curing, spar assembly, and wing assembly increase.

[0003] EP 2186622 A1 describes, according to its abstract, a composite structural member including first and second composite sections joined together by overlapping composite joining members. U.S. Patent No. 10836121 describes, according to its abstract, a method for fabricating a composite outer panel for a tiltrotor aircraft, the method including providing a first outer panel having a perimeter defined by a leading edge, a trailing edge, and an outboard end in a mold; providing a plurality of honeycomb panels having an array of large cells on the first outer panel, each cell having a width of at least 1 cm; assembling the plurality of honeycomb panels along the longitudinal axis of the first outer panel to form a honeycomb core having an outer perimeter within the perimeter of the first outer panel; positioning a second outer panel having an outer perimeter within the perimeter of the first outer panel on the honeycomb core; and curing an adhesive to create a bond between the first outer panel, the honeycomb core, and the second outer panel to form the composite outer panel. International Publication No. 2010 / 122325, according to its abstract, describes a structure comprising a cured composite component formed from a series of plies of fiber-reinforced composite material; a doubling plate attached to the composite component by an array of pointed prongs partially penetrating the composite component; and holes penetrating the doubling plate and the composite component. The interface plate carries the array of prongs on a first side and is attached to the doubling plate on a second side.

[0004] EP 3650333A1, according to its abstract, describes an embodiment of an aircraft wing comprising a first skin, a second skin opposite the first skin, and a composite spar. The composite spar includes a double-flange spar cap, a single-flange spar cap, a spar web connecting the double-flange spar cap and the single-flange spar cap, and a tear strap. The double-flange spar cap includes an inward-facing flange and a first outward-facing flange, which are integrated with the first skin during a co-curing process. The single-flange spar cap includes a second outward-facing flange attached to the second skin. The tear strap is sewn inside the spar web along at least a portion of the length of the composite spar. U.S. Patent Application Publication 2016 / 121589, according to its abstract, presents a method and related system for forming a composite material, the method comprising providing a composite material charge having a release film applied to its top and bottom surfaces, which is positioned on a contoured molding tool such that a portion of the composite material charge overlaps the second upper surface of the molding tool. A shaping bar contacts the molding tool having an uppermost surface positioned over the second upper surface of the molding tool. By combining the positioning of the shaping bar with the use of a release film in contact with the composite material charge, the composite material charge can be adapted to a contoured molding tool without bridging or wrinkling the composite material charge when film forming and vacuum application are used.

[0005] The aforementioned problem is further complicated by the fact that numerous spars exhibit complex curvatures corresponding to the curvature of the wing, which in turn complicates the manufacturing process.

[0006] Therefore, it would be desirable to have a method and apparatus that takes into account at least some of the above problems, as well as other possible problems. [Overview of the Initiative]

[0007] Embodiments described herein provide spars formed from multiple segments that are structurally joined together and integrated via co-curing, co-bonding, or fastener attachment. By assembling spars from multiple segments, the size of the equipment used to manufacture the spars can be reduced. Furthermore, since the segments are smaller than the entire spar, each segment exhibits a simpler curvature (e.g., a more linear profile) than the resulting spar, thereby further reducing the complexity of the equipment used for layup, curing, etc.

[0008] In one embodiment, a method for manufacturing spar detail for an aircraft is provided. The method includes manufacturing a preform of fiber-reinforced material for spar segments, solidifying the preform to form spar segments, and joining the spar segments together to form a finished spar detail.

[0009] In another embodiment, an aircraft spar is provided, comprising a first spar segment comprising a fiber-reinforced material, the first spar segment comprising a joining region, a second spar segment comprising a fiber-reinforced material, the second spar segment comprising the first joining region and arranged in series with the first spar segment, and a joining doubler covering at least a portion of the joining region of the first spar segment and the first joining region of the second spar segment, the joining doubler coupled to the first spar segment and the second spar segment.

[0010] In another embodiment, a method for manufacturing spar detail for an aircraft is provided. This method includes manufacturing fiber-reinforced preforms for a first spar segment and a second spar segment; joining the ends of the first spar segment preform and the ends of the second spar segment preform to define a joining area; applying at least one preform for joining doublers to the joining area; and simultaneously solidifying the spar segment and the preforms for joining doublers to form a portion of the spar detail.

[0011] In another embodiment, a spar detail is provided, comprising a preform for a first spar segment, wherein the first spar segment includes a bonding region; a preform for a second spar segment, wherein the second spar segment includes a bonding region; and a bonding doubling preform, wherein the bonding regions are arranged in series with respect to each other in a bonded relationship, and the preforms and bonding doubling preforms are cured simultaneously while the bonding doubling preform covers at least a portion of the bonded region to form a part of the spar detail.

[0012] In another embodiment, a method for manufacturing spar detailing for an aircraft is provided. This method includes manufacturing a preform of fiber-reinforced material for spar segments, solidifying the preform to form spar segments, and applying fasteners to join the spar segments together to form the finished spar detailing.

[0013] In yet another embodiment, a spar detail for an aircraft is provided. This spar detail comprises a first spar segment containing fiber-reinforced material, a second spar segment containing fiber-reinforced material and arranged in series with the first spar segment, a joining doubler covering a joining region between the first spar segment and the second spar segment, and fasteners attached through the joining doubler, the first spar segment, and the second spar segment to form at least a portion of the spar detail.

[0014] In another embodiment, a method for manufacturing an aircraft spar is provided. The method comprises manufacturing a preform of fiber-reinforced material of spar segments, wherein at least one of the spar segments includes a kink, and each kink is entirely contained within the preform; solidifying the preform to form spar segments; and assembling the spar segments together to form a finished spar detail exhibiting at least one of the kinks.

[0015] In another embodiment, an aircraft spar is provided, comprising a first spar segment comprising fiber-reinforced material, at least one bonding region, a second spar segment comprising fiber-reinforced material, at least one bonding region, and a kink outside the bonding region, wherein the respective bonding regions are arranged in series with respect to each other, and bonding doublers connecting the first and second spar segments, the bonding doublers covering at least a portion of the bonding region of the first spar segment and at least a portion of the corresponding bonding region of the second spar segment.

[0016] In another embodiment, a method for manufacturing spar detail for an aircraft is provided. The method includes manufacturing a preform for a first spar segment, wherein the preform includes a sub-kink adjacent to one end of the first spar segment; manufacturing a preform for a second spar segment, wherein the preform includes a sub-kink adjacent to one end of the second spar segment; aligning the ends of the preforms such that the sub-kinks are close to each other within a joining area; and joining the spar segments together within a joining area to form at least a portion of a spar detail exhibiting a kink.

[0017] In another embodiment, a spar detail for an aircraft is provided. The spar detail comprises a first spar segment comprising fiber-reinforced material and having a sub-kink located at one end; a second spar segment comprising fiber-reinforced material and having a sub-kink located at one end; and a joining doubling that structurally integrates the first spar segment and the second spar segment within a joining region, wherein the end of the first spar segment having the sub-kinks is adjacent to the end of the second spar segment having the sub-kinks such that the sub-kinks collectively form a kink and the end defines the joining region.

[0018] Other exemplary embodiments (e.g., methods and computer-readable media related to the embodiments described above) are also described later. The features, functions, and advantages described above can be realized individually or in combination in various embodiments, and further details of these embodiments can be found in the following description and drawings.

[0019] Herein, several embodiments of the present disclosure are described for illustrative purposes only with reference to the accompanying drawings. In all drawings, the same reference numerals represent the same element or element of the same type. [Brief explanation of the drawing]

[0020] [Figure 1] A perspective view of an aircraft including a fully assembled wing in an exemplary embodiment. [Figure 1A] A block diagram of a spar detail in an exemplary embodiment. [Figure 2] A system for forming a spar detail from a plurality of segments in an exemplary embodiment is shown. [Figure 3] A spar detail formed from a plurality of spar segments in an exemplary embodiment is shown. [Figure 4] A flowchart showing a method for manufacturing a spar detail from spar segments via co-bonding in an exemplary embodiment. [Figure 5] A flowchart showing a further method for manufacturing a spar detail from spar segments via co-curing in an exemplary embodiment. [Figure 6] A flowchart showing a method for manufacturing a spar detail from spar segments via a fastener in an exemplary embodiment. [Figure 7] A flowchart showing a method for manufacturing a spar detail having a kink in an exemplary embodiment. [Figure 8] A flowchart showing a further method for manufacturing a spar detail having a kink in an exemplary embodiment. [Figure 9] A wing of an aircraft including a spar in an exemplary embodiment is shown. [Figure 9A] A cross-section of a wing in an exemplary embodiment is shown. [Figure 10] A spar detail in an exemplary embodiment is shown. [Figure 11] A spar detail in an exemplary embodiment is shown. [Figure 12] A kink in a portion of a spar detail in an exemplary embodiment is shown. [Figure 13]This shows the assembly of a kink divided into smaller sections in an exemplary embodiment. [Figure 14] This shows the assembly of a kink divided into smaller sections in an exemplary embodiment. [Figure 15] This illustrates the formation of a flat charge onto a preform having a predetermined cross-section in one exemplary embodiment. [Figure 16] This illustrates the formation of a flat charge onto a preform having a predetermined cross-section in one exemplary embodiment. [Figure 17] This illustrates the formation of a flat charge onto a preform having a predetermined cross-section in one exemplary embodiment. [Figure 18] This shows a scarf joint between spar segments in an exemplary embodiment. [Figure 19] This is a flowchart illustrating a method for manufacturing and maintaining an aircraft in one exemplary embodiment. [Figure 20] This is a block diagram of an aircraft in one exemplary embodiment. [Modes for carrying out the invention]

[0021] The accompanying drawings and the following description provide specific exemplary embodiments of the present disclosure. Therefore, those skilled in the art will be able to devise various configurations that embody the principles of the present disclosure and fall within the scope of the present disclosure, although these are not expressly described or illustrated herein. Furthermore, any embodiments described herein should be interpreted as being for the purpose of understanding the principles of the present disclosure and not as limiting to such specifically enumerated embodiments and conditions. Consequently, the present disclosure is limited by the claims and their equivalents, and is not limited to the specific embodiments or examples described below.

[0022] The spars and spar details described herein may be manufactured as composite parts. Composite parts, such as carbon fiber reinforced polymer (CFRP) parts, are first laid up in a multi-layer structure called a preform. While the individual fibers within each layer of the preform are aligned parallel to one another, different layers exhibit different fiber orientations to increase the strength along different dimensions of the resulting composite part. The preform contains a viscous resin that solidifies to harden the preform into a composite part (for example, for use in aircraft). Uncured carbon fibers impregnated with thermosetting or thermoplastic resins are called "prepregs." Other types of carbon fibers include "dry fibers" that are not impregnated with thermosetting resins, but may contain tackifiers or binders. The dry fibers are injected with resin before hardening. In the case of thermosetting resins, hardening is a unidirectional process called curing, while in the case of thermoplastic resins, the resin can reach a viscous state when reheated, and then be compacted and solidified into the desired shape. As used herein, the collective term for the process of transitioning a preform to its final solidified form (i.e., transitioning a preform into a composite component) is "solidification," and this term encompasses both the curing of thermosetting preforms and the molding / solidification of thermoplastic preforms into their final desired shape.

[0023] Figure 1 illustrates a typical aircraft 10 on which exemplary embodiments of wing panels and / or wing assemblies manufactured according to one aspect of the present disclosure may be implemented. In other words, aircraft 10 is an embodiment of an aircraft that can be formed using composite components, wing panels and / or wing assemblies manufactured according to one or more aspects described herein. In this embodiment, aircraft 10 has wings 12 attached to and extending from either side of a fuselage 14. Aircraft 10 includes engines 16 attached to each wing 12. At the rear end of the fuselage 14 is a tail section 18 including opposing pairs of horizontal stabilizers 20 and vertical stabilizers 22. The wings 12 are formed of upper wing panels 30 and lower wing panels (not shown) joined together, with assemblies of ribs and spars (not shown) forming at least partially its internal structure.

[0024] Figure 1A is a block diagram of a spar detail 110 in an exemplary embodiment. As used herein, spar detail 110 refers to a structural component. Furthermore, spar detail may also be referred to as an aircraft spar detail or a multi-segment spar detail. Specifically, spar detail 110 is a component assembled from spar segments 112, 114, and 116, before being attached to other wing components such as ribs or wing panels as part of a finished or nearly finished product. Spar detail 110 becomes the rear spar 110-1 or front spar 110-2 (Figure 9) after other components (brackets, stiffeners, fasteners, etc.) are permanently attached. The “spar attachment section” (also referred to herein as “spar”) refers to the rear spar 110-1 and front spar 110-2 attached as part of the wing box (Figure 9). Extending from the inner end 197 to the outer end 199, spar detail 110 is one of the components that provide structural strength to the aircraft wing. For example, some spar details 110 extend from one side of the aircraft body intersection toward the wingtip 198. Some spar details 110 terminate at the wingtip 198, and in other embodiments terminate before reaching the wingtip 198. The spar details 110 can be thought of as the “core” of the spar (e.g., spars 110-1, 110-2 in Figure 9), providing most of the structural strength of the spar. Multiple longitudinal portions 127 of spar segments 112, 114, and 116 of the spar detail 110 are also shown in Figure 1A. For clarity, spar segment 112 is also referred to as the first spar segment 112, spar segment 114 as the second spar segment 114, and spar segment 116 as the third spar segment 116.

[0025] In this embodiment, the spar detail 110 is manufactured from a plurality of spar segments 112, 114, 116 (see also Figure 3), each of which defines one segment / part of the spar detail 110 from the machine-side end 197 to the machine-side end 199. Each spar segment 112, 114, 116 includes multiple layers 122 of fibers 124 (e.g., continuous carbon fiber, glass fiber, etc.) and resin 126 (e.g., thermosetting resin, thermoplastic, etc.). The spar detail 110 is manufactured from spar segments 112, 114, 116 which are pre-molded and / or solidified before assembly. Thus, the entire manufacturing process of spars 110-1, 110-2 manufactured from spar segments 112, 114, 116 can be carried out quickly and efficiently, for example, from each parallel spar segment 112, 114, 116, using less expensive and less bulky equipment than a manufacturing system that does not manufacture spars from segments.

[0026] Figure 2 shows a system 200 for forming spur detail 110 from spur segments 112, 114, and 116 in an exemplary embodiment, for example, shown in Figure 1. In this embodiment, the system 200 includes a placement machine 230 that distributes tows of fiber-reinforced material to form, for example, one or more single-ply flat charges 232 (e.g., planar charges) or, when combined with other flat charges 232, to form a complete flat charge 233 laminate on a mandrel 234. In this embodiment, the flat charges 232 and / or complete flat charges 233 are formed by a molding machine 240 to produce a preform 242. The preform 242 is then placed on a curing mandrel 238 by an end effector 262 of a pick-and-place (PNP) machine 260.

[0027] The preform 242 is a laminate of multiple flat charges 232 and / or complete flat charges 233 that have been molded and assembled together. One or more of the flat charges 232 proceed to the molding machine 240 for ply-by-ply (PBP) molding and placement as part of the preform 242. Shaping or forming of the flat charges 232 or complete flat charges 233 is carried out by draping, stamping, ply-by-ply molding or other suitable forming method by the molding machine 240. Manufacturing the preform 242 includes distributing tows of fiber-reinforced material to form the flat charges 232, assembling multiple flat charges 232 into a laminate 235 to form a complete flat charge 233, and shaping the complete flat charge 233 into a preform 242 having the desired cross-sectional shape. The molding machine 240 shapes a flat charge 232, which is placed on a curing mandrel 238, as part of producing a preform 242 that will solidify into spur segments 112, 114, and 116 in an autoclave 250 after the application of a vacuum bag 252 and, if possible, a backing plate 253. Another alternative example has a combined vacuum bag and a backing plate (not shown). The preform for the bonding doubling 340 (see Figure 3) is supplied from a bonding doubling feeder line 244. The preform for the bonding doubling 340 may also be called a bonding doubling preform.

[0028] In one embodiment not shown, a vacuum bag 252 covers preforms 242 for a plurality of spar segments 112, 114, and 116, and the entire assembly is co-cured. In this way, the preforms 242 are solidified together with the preforms for the joining doublers 340 for integration with the ribs 290 and the wing skin 30, and solidified in an autoclave 250 into the spar detail 110 shown in Figure 1. The preforms for the joining doublers 340 are assembled for just-in-time (JIT) delivery for assembly together with the preforms 242.

[0029] As shown in the figure, the spur segments 112, 114, and 116 are solidified from the preform 242 in an autoclave 250. The spur segments 112, 114, and 116 are then aligned end to end with each other. The spur segments 112, 114, and 116 are combined with a joining doubler 340 and assembled together using fasteners to form the spur detail 110. The spur segments 112, 114, and 116 are then assembled together via a segment splicer 280 in the form of operation of a fastener mounting station 272. In this embodiment, the fastener mounting station 272 includes a jig 276 that holds the spur segments 112, 114, and 116 and the solidified joining doubler 340 relative to each other, and an end effector 274 for mounting fasteners 278 through these components. Specifically, the end effector 274 drills mounting holes and drives a fastener 278 (e.g., a lock bolt) to join the bonding doubling 340 to the spur segments 112, 114, and 116. In this embodiment, the fastener 278 is driven through the bonding doubling 340 and the spur segments 112, 114, and 116 to form a single spur detail 110. The solidified bonding doubling 340 is assembled for just-in-time delivery for assembly with the multiple spur segments 112, 114, and 116.

[0030] The preform 242 is solidified separately into spur segments 112, 114, and 116 via autoclave 250, and then aligned end to end (in series). The end effector 271 positions the preform for the bonding doublers 340 in the bonding regions 341, 341-1 (Figure 1A) between the spur segments 112, 114, and 116. The pressclave 270 is then used on at least a portion of the segment splicer 280 to co-bond the preform for the bonding doublers 340 to the spur segments 112, 114, and 116 in order to form the completed spur detail 110. The pressclave 270 solidifies one or more of the bonding doublers 340 in place at the intersections between the spur segments 112, 114, and 116 through the application of heat and pressure.

[0031] Another alternative is the combination of the pressclave 270 process and spar segments 112, 114, and 116 joined by fasteners 278. The pressclave 270 process would precede the fastener attachment process.

[0032] The operation of system 200 is managed by server 220, which includes memory 222 and controller 224. In one embodiment, controller 224 is implemented as a custom circuit, as a hardware process that executes programmed instructions stored in memory 222, or as any combination thereof.

[0033] System 200 can simultaneously manufacture multiple sets of spar segments 112, 114, and 116 for multiple spar details 110. System 200 is divided into multiple manufacturing hemispheres, each capable of independently manufacturing a set of spar segments 112, 114, and 116 that will ultimately be assembled into a spar detail 110. Therefore, the solidification of the preforms 242 for the spar segments 112, 114, and 116 is carried out while the upstream preforms 242 are assembled together from the flat charge 232, the complete flat charge 233, and the molding machine 240. This ensures that the assembly of the spar segments 112, 114, and 116 is carried out simultaneously (in parallel) with the solidification of the upstream sets of spar segments 112, 114, and 116 in the autoclave 250. Post-solidification joints of further upstream sets of spar segments 112, 114, and 116 by the segment splicer 280 are formed in another spar detail 110. The various upstream and downstream spur segments 112, 114, and 116 typically consist of a front spur 110-2 and a rear spur 110-1 and / or a right spur and a left spur.

[0034] Figure 3 shows a spar detail 110 (having a width W that gradually decreases from the inner end 197 to the outer end 199 and a length L) formed from a first spar segment 112, a second spar segment 114, and a third spar segment 116 in an exemplary embodiment. The first spar segment 112 contains fiber reinforced material 301, and the second spar segment 114 contains fiber reinforced material 301-1 and is arranged in series with the first spar segment 112. The first spar segment 112 has an inner end 311 and an outer end 312. The outer end 312 of the first spar segment 112 is arranged end-to-end with the inner end 321 of the second spar segment 114. The second spar segment 116 has an inner end 321 and an outer end 322. The joining doubler 340 covers a joining region 341 (also known as the “intersection” or “joining zone”) including the machine outer end 312 and the machine inner end 321, and is located in the joining region 341 between the first spur segment 112 and the second spur segment 114. Depending on the embodiment, the joining region 341 may be implemented as a butt splice, scarf splice, wrap splice, or other splice of the web and flange of the machine outer end 312 of the first spur segment 112 and the machine inner end 321 of the second spur segment 114 with the joining doubler 340 to complete the joining. Similarly, the third spur segment 116 includes fiber-reinforced material 301-2 and is located in series with the second spur segment 114. The third spur segment 116 has a machine inner end 331 and a machine outer end 332. The outer end 322 of the second spur segment 114 is aligned end-to-end with the inner end 331 of the third spur segment 116. Another bonding doubling 340-1 covers the bonding region 341-1 between the outer end 322 of the second spur segment 114 and the inner end 331 of the third spur segment 116. In one embodiment, an adhesive 319 (e.g., epoxy, thermosetting resin, etc.) is used to bond the bonding doubling 340, the first spur segment 112, and the second spur segment 114 to a single composite component.Depending on the embodiment, the joining region 341 may be implemented as a butt splice, scarf splice, lap splice, or other splice of the web and flange of the machine-side end 322 of the second spur segment 114 and the machine-side end 331 of the third spur segment 116, with a joining doubler 340-1 added to complete the joining. In such embodiments, the adhesive 319 can also bond the other joining doubler 340-1, the second spur segment 114, and the third spur segment 116 to the spur detail 110. In further embodiments, these components may be fastened together, fastened and joined, co-cured, or joined together.

[0035] In Figure 3, one of the spar segments (specifically the second spar segment 114) completely contains the kink 370, but in further embodiments, the kink 370 is distributed across the joint region 341, and multiple kinks (multiple kinks are not shown) are each completely contained within spar segment 114, or each of multiple spar segments, such as spar segment 112, spar segment 114, and spar segment 116, completely contains the kink 370. Each kink 370 contains a bending point 381, at which the intersection of the first neutral axis 350 and the second neutral axis 350-1 of the spar detail 110 is located. In one embodiment, the variation is shown by a bending angle θ of 2 to 10 degrees. In one embodiment, the kink 370 includes a variation in the axial direction of the spar detail 110, which is, for example, 2 to 10 degrees. In some embodiments, as shown in Figure 3, the kink 370 is located 380 from the machine-side neutral axis endpoint 322 (machine-side end 322) and 390 from the machine-side neutral axis endpoint (machine-side end 321) of a spar segment, for example, a second spar segment 114, more than approximately 30 cm (1 foot). In further embodiments, the flat charge 232 (or the resulting preform 242) in Figure 2 is prepared to fully include the kink 370 before solidifying into the spar segment 114. In one embodiment, to increase the structural strength of the spar detail 110, the kink 370 is located in a joint region 341, and a rib 290 (not shown) is also located in the joint region 341. The first spar segment 112 may include the kink 370 outside the joint region 341, for example.

[0036] Each spur segment 112, 114, and 116 exhibits a different shape to accommodate the reduction in the thickness and / or bending angle θ of the kink 370 due to design parameters. Figure 3 further illustrates that the joining doubler 340 facilitates the incorporation of the spur segments 112, 114, and 116 into the spur detail 110. The rib intersection 360 is positioned in and / or between the joining doublers 340, 340-1 to receive the rib 290 that further supports the joining region 341, 341-1. The rib intersection 360 is shown as a small rectangular box, but this does not imply that the complex connection of the rib 290 to the spur detail 110 is limited to this relatively small area. In one embodiment, the rib intersection 360 is positioned in the spur detail 110 opposite the joining doubler 340, shown on the opposite side 391 from the side to which the joining doubler 340 is attached. Thus, the joining doubler 340 is attached to the side of the spur detail 110 that is not visible in this figure, while the rib 290 is attached to the rib intersection 360 on the visible side 391 of the spur detail 110. The joining doubler 340 can be positioned across the rib intersection 360.

[0037] The spar detail 110, assembled from individual spar segments 112, 114, and 116, offers substantial advantages over prior art implementations by reducing the size and complexity of the equipment required to manufacture the spar detail 110. Furthermore, the spar segments 112, 114, and 116 are advantageously manufactured in parallel, thus substantially reducing manufacturing time and increasing work density. This results in technical advantages such as cost reduction and space savings on the factory floor, which in turn increases work density. Along with the above-described spar detail 110 and the system for manufacturing it, Figure 4-8 below illustrates various methods for manufacturing the spar detail 110.

[0038] Exemplary details of the operation of the joint 200 are described with reference to Figure 4. In this embodiment, it is assumed that the components of the system 200 are awaiting operation to manufacture spar segments that will be assembled into one or more spars 110.

[0039] Figure 4 is a flowchart illustrating a method 400 for manufacturing spar detail 110 from spar segments 112, 114, and 116 via cobonding in an exemplary embodiment. The steps of method 400 are described with reference to system 200 in Figure 2, but those skilled in the art will understand that method 400 may be carried out in other systems. The steps in the flowcharts described herein are not exhaustive and may include other steps not shown. The steps described herein may be carried out in a different order.

[0040] The placement machine 230, together with the molding machine 240, produces a preform 242 of fiber-reinforced material (e.g., CFRP) for spar segments 112, 114, and 116 402. In one embodiment, producing the preform 242 402 includes the placement machine 230 laying up one or more flat charges 232. The flat charges 232 are used either directly to produce the preform 242 or are combined into a complete flat charge 233 before being used to produce the preform 242. The flat charges 232 and / or the complete flat charge 233 are then molded by the molding machine 240 and then placed in the preform 242 having the desired cross-section. In further embodiments not shown, producing the preform 242 includes laying up a tow of fiber-reinforced material on a curing mandrel 238 that defines the shape of the preform 242. In further embodiments, the production of the preform 242 includes distributing tows of fiber-reinforced material to form flat charges 232, and forming the flat charges 232 before placement during the production of the preform 242 having a desired cross-sectional shape.

[0041] The preform 242 is solidified 404 to form spar segments 112, 114, and 116. In one embodiment, solidifying the preform 242 404 includes compacting and solidifying a thermoplastic resin, while in a further embodiment, solidifying 404 includes heating a thermosetting resin to a curing temperature while under compaction pressure in an autoclave 250. In any embodiment, the resulting spar segments 112, 114, and 116 contain solidified fiber-reinforced material shaped according to design parameters for specific longitudinal portions 127 of the spar detail 110.

[0042] Spur segments 112, 114, and 116 are joined together 406, with the outer end 312 abutting or joining to the inner end 321 and the outer end 322 abutting or joining to the inner end 331 to form a spur detail 110. In one embodiment, this includes applying adhesive 319 or resin to the spur segments 112, 114, and 116 and forming a joint between the spur segments 110 within a joining region 341 (see Figure 3). In a further embodiment, this includes placing an unsolidified "green" joining doubling 340 within the joining region 341 between spur segments 114 and 112 and solidifying the joining doubling 340 via cobonding within a pressclave 270. This operation integrates the joining doubling 340 with the spur segments 114 and 112, forming a joined first spur segment 112 and second spur segment 114. Thus, in one embodiment, joining the spar segments 112 and 114 together 406 includes applying a joining doubler 340 that partially covers the first spar segment 112 and the second spar segment 114, followed by joining the joining doubler 340 to the first spar segment 112 and the second spar segment 114 406. In some embodiments, the joining doubler 340 is positioned so that it extends across the rib intersection 360 of the spar detail 110 in order to reinforce the joint. Thus, in one embodiment, applying the joining doubler 340 includes sandwiching the first spar segment 112 and the second spar segment 114 between a front joining doubler 340 and a rear joining doubler (not shown). In such embodiments, the joining doubling 340 includes either a front joining doubling or a rear joining doubling, and the spar includes the other of the front joining doubling or the rear joining doubling, with the joining region 341 of the first spar segment 112 and the first joining region 341 of the second spar segment 114 sandwiched between the front joining doubling and the rear joining doubling. In some embodiments, the joining doubling 340 sandwiches a portion of the first spar segment 112 and a portion of the second spar segment 114 between the joining doubling 340 and either the second joining doubling or the rib 290.

[0043] In another embodiment, the application of the joining doubler 340 sandwiches the first spur segment 112 and the second spur segment 114 between the joining doubler 340 and the rib 290. In such an embodiment, the joining region of the first spur segment and the first joining region of the second spur segment are sandwiched between the joining doubler and the rib. Similarly, the second spur segment 114 and the third spur segment 116 are joined together using the joining doubler 340 and / or the rib 290 in a manner similar to the above-described method used to join the first spur segment 112 to the second spur segment 114, resulting in the spur detail 110.

[0044] After completion, the spar detail 110 is assembled together with further details to form a finished spar, such as the rear spar 110-1 and front spar 110-2 in Figure 9. Method 400 offers substantial advantages over prior art systems because the spar detail 110 can be formed via co-bonding from a series of smaller individual spar segments 112, 114, and 116, each occupying a portion of the overall length of the spar detail 110. This reduces manufacturing complexity and shortens manufacturing time. Furthermore, if rework for individual spar segments 112, 114, and 116 is guaranteed, rework can be performed on individual spar segments 112, 114, and 116 instead of the entire spar detail 110.

[0045] Figure 5 is a flowchart illustrating a further method 500 for manufacturing spur detail 110 from spur segments 112, 114, and 116 via co-curing in an exemplary embodiment. Method 500 manufactures a fiber-reinforced material preform 242 for the spur segments 112, 114, and 116 502, which can be achieved in a manner similar to the manufacturing steps described above. For example, prior to co-curing, the first spur segment 112, the second spur segment 114, the third spur segment 116, the bonding doubling 340, and the second bonding doubling 340-1 may contain dried fibers injected in place of resin before co-curing. As part of this manufacturing process, the preform 242 may include ramps, stepped patterns, and scarf-type bonding of complementary features of the preform 242 within a bonding region 341. Within the bonding region 341, the machine-outer end 312 is butted against or bonded to the machine-inner end 321. Within the joining region 341-1, the machine outer end 322 is abutted against or joined to the machine inner end 331. In one embodiment, manufacturing the preform 242 502 includes distributing / laying up tows of fiber-reinforced material to form flat charges 232, and shaping the flat charges 232 into a preform 242 having a desired cross-sectional shape. In a further embodiment, laying up the preform 242 includes combining multiple flat charges 232 into a complete flat charge 233, and shaping the complete flat charge 233 into a preform 242 having a desired cross-sectional shape.

[0046] The preforms 242 for spar segments 112 and 114 are joined together 504 with the machine-side end 312 of the first spar segment 112 aligned end to end with the machine-side end 331 of the spar segment 114. In other words, the spar segments 112 and 114 are arranged in series. In one embodiment, joining the preforms 242 for spar segments 112 and 114 504 is carried out within a joining region 341 by a lap splice, butt splice, and / or scarf splice of the complementary ramp or pattern of the preform. The preforms 242 for spar segments 114 and 116 are joined together 504 with the machine-side end 322 of the first spar segment 112 aligned end to end 331 of the spar segment 116. In one embodiment, joining the preforms 242 for spar segments 114 and 116 504 is carried out by a wrap splice, butt splice, and / or scarf splice of the complementary ramp or pattern of the preform 242 within the joining region 341-1. In a further embodiment, joining the preforms 242 for spar segments 114 and 116 504 is carried out by arranging a dry fiber preform together and then injecting resin into the preform 242.

[0047] Next, the PNP machine 260 applies the preform for the joining doubler 340 to the joining region between the preforms 242 for the spar segments 112 and 114 506. In one embodiment, this operation includes placing the unsolidified preform for the joining doubler 340 in the joining region 341 that forms the joint between the preforms 242 for the spar segments 112 and 114. In another embodiment, applying the preform for the joining doubler 340 506 includes picking up the preform for the joining doubler 340 and positioning it so that it extends across the rib intersection 360 of the spar detail 110. In a further embodiment, applying the preform for the joining doubler 340 506 includes laying up the preform for the joining doubler 340 (not shown) so that it extends across the rib intersection 360 of the spar detail 110. In one embodiment, applying a preform (not shown) for a joining doubler 340 includes laying up a preform for the joining doubler 340 that extends across the outer end 312 and the inner end 321 of the machine, and across from the rib intersection 360 of the spar detail 110 to the opposite side of the spar detail 110. In this way, the rib intersection 360 is located on the opposite side of the joining doubler 340 of the spar detail 110. In addition, applying a preform (not shown) for a joining doubler 340-1 includes laying up a preform for the joining doubler 340-1 that extends across the outer end 312 and the inner end 321 of the machine, and is located on the opposite side of the spar detail 110 from the rib intersection 360 of the spar detail 110. Similarly, the rib intersection 360 is located on the opposite side of the joining doubler 340-1 of the spar detail 110.

[0048] Autoclave 250 solidifies preforms 242 for spur segments 112, 114, and 116 and preforms for bonding doublings 340, 340-1 between 504 where the preforms for spur segments 112, 114, and 116 are joined together to form the completed spur detail 110 508. In a further embodiment, solidifying preforms 242 for spur segments 112, 114, and 116 and preforms for bonding doublings 340, 340-1 508 includes vacuum bagging of preforms 242 for spur segments 112, 114, and 116 and preforms for bonding doublings 340, 340-1. The preforms 242 for spur segments 112, 114, and 116 and the preforms for bonding doublings 340, 340-1 are compacted by vacuum compression and heating the preforms 242 for spur segments 112, 114, and 116 and the preforms for bonding doublings 340 to a curing temperature or compaction temperature. Thus, in one embodiment, the vacuum bag 252 further covers the preform 242 for the third spur segment 116 and the preforms for the second bonding doublings 340, 340-1.

[0049] Figure 6 is a flowchart showing a further method 600 for manufacturing spur detail 110 from spur segments 112, 114, and 116 via fasteners in an exemplary embodiment. A placement machine 230 and / or a molding machine 240 manufactures a preform 242 of fiber-reinforced material for the spur segments 112, 114, and 116 602. Manufacturing 602 may be carried out in a manner similar to the manufacturing step 402 in Figure 4 described above. In one embodiment, manufacturing the preform 242 602 includes distributing a tow of fiber-reinforced material to form a flat charge 232 and shaping the flat charge 232 into a preform 242 having a desired cross-sectional shape. In a further embodiment, manufacturing the preform 242 602 includes distributing a tow of fiber-reinforced material to form a flat charge 232, combining a plurality of flat charges 232 into a laminate 235 to form a complete flat charge 233, and shaping the complete flat charge 233 into a preform 242 having a desired cross-sectional shape. Embodiments may include the use of dry fibers and resin injection, as described in other parts of this specification.

[0050] The preform 242 is solidified to form spur segments 112, 114, and 116.604 In one embodiment, this involves operating an autoclave 250 to solidify the preform 242 into individual spur segments 112, 114, and 116.604 At this point, the spur segments 112, 114, and 116 remain physically separated from one another. In a further embodiment, joining doublings 340, 340-1 are used to form joints and fastened to each spur segment 112, 114, and 116.

[0051] The jig 276 of the fastener mounting station 272 holds the spur segments 112 and 114, and additionally 114 and 116, and the solidified joining doublers 340, 340-1 in joining regions 341, 341-1, respectively. The end effector 274 mounts fasteners that connect these components to the spur detail 110. Specifically, the end effector 274 applies fasteners 606 that connect spur segments 112 to 114 and 114 to 116 together to form the completed spur detail 110. In one embodiment, applying fasteners 606 includes applying joining doublers 340, 340-1 that partially cover the machine outer end 312 and the machine inner end 321, and mounting fasteners through the joining doublers 340 and through the first spur segment 112 and the second spur segment 114. 606 applies fasteners to a joining doubler 340-1 that partially covers the outer end 322 and the inner end 331 of the machine, and fastens through the joining doubler 340-1 and through the first spur segment 112 and the second spur segment 114. In a further embodiment, applying the joining doubler 340 includes positioning the joining doubler 340 in the joining areas 341, 341-1 of the spur detail 110 so that the joining doubler 340 extends across the rib intersection 360. In one embodiment, applying fasteners for the joining doubler 340 includes positioning the joining doubler 340 to span the outer end 312 and the inner end 321, and to extend from the rib intersection 360 of the spur detail 110 to the opposite side of the spur detail 110. In this way, the rib intersection 360 is positioned on the opposite side of the joining doubler 340 of the spur detail 110. After assembly, the spur detail 110 is combined with other details to form spurs 110-1 and 110-2 (Figure 9). In this way, the fastener mounting station 272 assembles the joining doublers 340 and 340-1 and the spur segments 112, 114, and 116 into a single, integrated spur detail 110.

[0052] Figure 7 is a flowchart showing a method 700 for manufacturing a spar detail 110 having one or more kinks 370 in an exemplary embodiment. The method 700 includes laying up or manufacturing 702 a preform 242 of fiber-reinforced material to be assembled into a spar detail 110 containing one or more kinks 370. Each kink 370 is entirely contained within a single preform 242 for spar segments 112, 114, and 116, for example. In one embodiment, manufacturing 702 the preform 242 includes changing the orientation of each kink 370 from a first neutral axis 350 to a second neutral axis 350-1 of the preform 242 (for example, during layup or molding), the change being between 2 and 10 degrees. In further embodiments, manufacturing the preform 242 702 includes positioning each kink 370 at least 30 cm (1 foot) from the machine-external end 322 and machine-internal end 321 of the preform 242 for the spur segment 114. Similarly, positioning the kink 370 within the spur segment 112 and / or spur segment 114 is also considered. In further embodiments, laying up the preform 242 (i.e., manufacturing 702) includes laying up the flat charge 232 on the mandrel 234, followed by shaping the flat charge 232 on a rigid tool, such as a hardened mandrel 238. Shaping or forming the flat charge 232 or the complete flat charge 233 is carried out by drape forming, stamp forming, ply-by-ply forming or other suitable forming method before incorporation into the preform 242.

[0053] The preform 242 is solidified to form spur segments 112, 114, and 116 704, and the spur segments 112, 114, and 116 are then assembled together to form a finished spur detail 110 exhibiting a kink 370 706. The assembly process may be carried out as described above, via co-curing, co-bonding, or via fasteners or combinations of fasteners and bonding or co-bonding. In one embodiment, the solidification 704 of the preforms 242 for spur segments 112, 114, and 116 is carried out while the upstream preforms 242 are assembled together from the flat charge 232, the complete flat charge 233, and the molding machine 240, and the assembly of the spur segments 112, 114, and 116 is carried out simultaneously with the solidification 704 of the upstream set of spur segments 112, 114, and 116 in the autoclave 250, and the post-solidification bonding of the further upstream set of spur segments 112, 114, and 116 to the spur detail 110 by the segment splicer 280. The upstream and downstream spur segments typically consist of a front spur 110-2 and a rear spur 110-1 and / or a right spur and a left spur set.

[0054] The spur segments 112 and 114 are assembled together to form a portion of the finished spur detail 110 exhibiting a kink 370 706. In one embodiment, assembling the spur segments 112 and 114 706 includes applying a joining doubler 340 across the outer end 312 and the inner end 321 to join the spur segment 112 to the spur segment 114. Assembling the spur segments 114 and 116 706 includes applying a joining doubler 340-1 across the outer end 322 and the inner end 331 to join the spur segment 114 to the spur segment 116. In some embodiments, assembling the spur segments 112, 114, 116 706 includes forming at least one of lap splices, butt splices and scarf splices between adjacent ends of the spur segments 112, 114, 116, and attaching bonding doublings 340, 340-1 to the bonding regions 341, 341-1 using one of co-curing, co-bonding and fastener. In some embodiments, assembling the spur segments 112, 114, 116 706 includes aligning the spur segments 112, 114, 116 in series with each other to form the bonding regions 341, 341-1, and applying bonding doublings 340, 340-1 to the spur segments 112, 114, 116 within the bonding regions 341, 341-1.

[0055] In certain applications, a single kink 370 can be implemented, while in other applications, a single kink 370 may generate an angle between two neutral axes that are too large to be manufactured using known CFRP manufacturing methods. For this purpose, Figure 8 is a flowchart illustrating a further method 800 for manufacturing the spur detail 110. In the embodiments shown in method 800, multiple kinks are incorporated relatively close to each other across adjacent spur segments (e.g., spur segments 112 and 114) to allow for a smoother transition of the change in the neutral axis between adjacent spur segments 112, 114. In such embodiments, for clarity, the multiple kinks are referred to herein as sub-kinks 371, 371-1, and are shown in Figure 13.

[0056] In some embodiments, manufacturing the preform 242 702 includes changing the axial direction of the preform 242 at each kink 370. In some embodiments, manufacturing the preform 242 702 includes changing the axial direction of the preform 242 at each kink 370, where the axial change is a bending angle of 2 to 10 degrees. In some embodiments, each preform 242 includes a first end and a second end opposite the first end, and manufacturing the preform 242 702 includes positioning each kink 370 at least 1 foot from the end of the preform 242.

[0057] As shown in Figure 8, the fiber-reinforced material preform 242 is manufactured into spar segments 112, 114, and 116, which are assembled into a spar detail 110 containing a number of subkinks 371, 371-1 802. For example, the ends of the boundary between spar segment 112 and spar segment 114 each incorporate subkinks, e.g., 371, 371-1, and each subkink 371-371-1 changes the orientation of the preform 242 of the spar detail 110 from the first neutral axis 350 to the second neutral axis 350-1. Manufacturing the preform 242 802 can be carried out as described above in the aforementioned method.

[0058] The ends of the preform 242 having subkinks 371, 371-1 (e.g., spar segments 112, 114) are aligned within the joining area so that the subkinks 371, 371-1 are close to each other 804. This will be described further later with reference to Figure 13. The spar segments 112 and 114 are then joined together within the joining area 806. Aligning the ends 804 and joining 806 includes overlapping, butt joining, scarf joining, or other means during the production of the preform 242.

[0059] The spur detail 110 is manufactured using the preform 242. In one embodiment, the manufacturing of the spur detail 110 is carried out by co-curing the preform 242, co-bonding the spur segments 112, 114, and 116, fastening the spur segments 112, 114, and 116, or by co-bonding and fastening the spur segments 112, 114, and 116 together, as described above.

[0060] Figure 9 shows an aircraft wing 900 including spars in an exemplary embodiment. Figure 9A shows a cross-section of wing 900 (similar to wing 12 in Figure 1) corresponding to arrow 9A in Figure 9. Wing 900 is an exemplary embodiment of an aircraft including a rear spar 110-1 and a front spar 110-2 in an exemplary embodiment. The rear spar 110-1 and the front spar 110-2 are concealed beneath the wing panel 75 (similar to panel 30 in Figure 1), and each extends from the inner end 197 to the outer end 199 and the wingtip 198, although in embodiments not shown, some spar details 110 may terminate before reaching the wingtip 198. For example, some spar details 110 extend from one side of the aircraft body intersection toward the wingtip 198. The front spar 110-2 and the rear spar 110-1 terminate at the wingtip 198. As shown in Figure 9A, the rear spar 110-1 includes a first flange 902 and a second flange 906, with a web 904 between them. Briefly, the front spar 110-2 also includes a first flange 902, a second flange 906, and a web 904.

[0061] Typical wing panels 75, 76 include a composite wing skin and composite stringer attached to the wing skin and extending from the inner end 197 to the outer end 199. As will be described later, the wing panels 75, 76 are formed of a composite material such as CFRP, which is laid up to band widths of strips having various fiber orientations to provide the desired strength and flexibility. As shown in Figure 9, the ribs 290 are aligned along the chord direction below the wing panel 75, between the rear spar 110-1 and the front spar 110-2.

[0062] Figure 9A shows the front spar 110-2 and rear spar 110-1 positioned between wing panels 75 and 76 and coupled to rib 290. To complete the wing 900, wing panel 75 is coupled to flange 906 and rib 290, respectively, and wing panel 76 is coupled to flange 902 and rib 290, respectively. The illustrated wing panels 75 and 76 do not have stringers for clarity.

[0063] Figures 10-11 show a spar detail 110 in an exemplary embodiment. This represents either the rear spar 110-1 and / or the front spar 110-2. Figure 10 shows a first view in which the web 904 and flange 902 are visible, and Figure 11 shows a second view corresponding to arrow 11 in Figure 10. In Figure 11, each flange 902 and 906 is clearly visible, as is the web 904 connecting to the flanges 902 and 906. Each kink 370 includes a bending point 381 (see Figure 12) where the intersection of the first neutral axis 350 and the second neutral axis 350-1 of the spar detail 110 is located, all of which are shown in Figure 12.

[0064] Figure 12 shows a kink 370 in a portion of a spar detail 110 in an exemplary embodiment, corresponding to the figure in the box labeled Figure 12 in Figure 3. The kink 370 is the result of the first neutral axis 350 (shown by the dotted line) of the spar detail 110 bending at a bending point 381 by an angle θ to become the second neutral axis 350-1. Integrating the kink 370 into the spar detail 110, which is formed from multiple segments (not shown in Figure 12), can complicate the assembly process. Therefore, in one embodiment, the kink 370 is entirely contained within a single spar segment 112, 114, and 116. As shown in Figure 12, the kink 370 is entirely contained within the second spar segment 114. Since the kink 370 is entirely contained within the spur segments 112, 114, and 116, the different spur segments 112, 114, and 116 of the spur detail 110 are joined in joining regions 341 and / or 341-1 (both see Figure 3), which exhibit a neutral axis aligned with the first neutral axis 350 and / or the second neutral axis 350-1, respectively. Such alignment of the neutral axis with respect to the first neutral axis 350 and / or the second neutral axis 350-1 substantially reduces the difficulty of aligning and mounting the spur segments 112, 114, and 116 together.

[0065] In a further embodiment, as shown in Figures 13 and 14, the kink 370 having an angle θ is subdivided into two sub-kinks 371 and 371-1 by bends 1312 and 1322 of adjacent spar segments, e.g., spar segments 112 and 114, having angular deviations of θ2 and θ3, respectively. At bend 1322, the first neutral axis 1324 transitions to the second neutral axis 1324-1 by an angle of θ2. Sub-kink 371-1 is the result of the first neutral axis 1324 of the spar detail 110 bending at the bend point 381-1 by an angle of θ2 to become the second neutral axis 1324-1. At bend 1312, the second neutral axis 1324-1 transitions to the third neutral axis 1314 by an angle of θ3. The sub-kink 371 is the result of the second neutral axis 1324-1 of the spur detail 110 bending at the bending point 381-2 by an angle θ3 to become the third neutral axis 1314. In one embodiment, the sum of the angle θ2 of sub-kink 371-1 and the angle θ3 of sub-kink 371 is equal to angle θ, which may or may not be equal to the angle θ of the single kink 370 configuration shown in Figure 12.

[0066] In a further embodiment, the subkinks 371, 371-1 are distributed among further spar segments 112, 114, 116. For example, the intersection of spar segments 112 and 114 includes two subkinks 371, 371-1, and the intersection of spar segments 114 and 116 also includes two subkinks 371, 371-1.

[0067] Thus, when the spar segments 112, 114, and 116 are joined, the assembly process can be accurately carried out and a secure bond can be ensured by aligning the spar segments 112, 114, and 116 together 804 and applying the joining doubler 340 (Figure 3). The joining doubler 340 can co-harden, co-bond, or fasten to any two of the spar segments 112, 114, and 116.

[0068] To reiterate, in further embodiments, a first spar segment, e.g., spar segment 112, comprises fiber-reinforced material and includes a sub-kink 371-1 located at its end 1310, and a second spar segment, e.g., spar segment 114, also comprises fiber-reinforced material and includes another sub-kink 371 located at its end 1320. In some embodiments, the sub-kink 371-1 associated with the first spar segment 112 is located within approximately 30 cm (1 foot) of the end of the first spar segment 112; the sub-kink 371 associated with the second spar segment 114 is located within approximately 30 cm (1 foot) of the end of the second spar segment 114. The end 1310 is located adjacent to the end 1320 (e.g., within 30 cm / 1 foot of the end), and the spar segments 112, 114 in this embodiment are arranged in series such that the sub-kinks 371, 371-1 together provide a desired amount of kink. A component (for example, the joining doubler 340 in Figure 3 above) can structurally join the first spar segment 112 and the second spar segment 114. In some embodiments, the joining doubler 340 is positioned so as to sandwich the joining region 341 and sub-kinks 371, 371-1 between the joining doubler 340 and the second joining doubler. In some embodiments, manufacturing the preform 242 802 includes manufacturing the preform 242 such that each sub-kink 371-1, 371 has equal angular deviations. In some embodiments, manufacturing the preform 242 802 includes manufacturing the preform 242 such that the sub-kinks 371-1, 371 have unequal angular deviations. In some embodiments, each sub-kink 371-1, 371 changes the axial direction of the spar detail 110 by half the amount of the kink 370.

[0069] In the embodiment shown in Figure 14, a spar segment, for example, spar segment 116, exhibits a single bend 1312 of a different size from another bend 1322 seen in spar segment 114. The sum of the angle θ2 of sub-kink 371-1 and the angle θ3 of sub-kink 371 is still equal to angle θ, but angle θ2 is not equal to angle θ3.

[0070] In some embodiments, subkinks 371, 371-1 are formed on the bent portions 1312 and 1322 of a flat charge 232, which includes bent portions 1312 and 1322 in the layup. Another alternative example has a molding machine that shapes a flat charge 232 without subkinks 371-, 371-1 into a preform 242 which includes subkinks 371, 371-1 and bent portions 1312 and 1322 at ends 1310 and 1320 that combine to form a kink of angle θ.

[0071] Figures 15-17 show the formation of a flat charge 232 into a preform 242 having a predetermined cross-section 1702 in an exemplary embodiment. In Figure 15, one or more layers of flat charge 232 are formed by the application of multiple tows 1510. In Figure 16, corresponding to arrow 16 in Figure 15, a complete flat charge containing multiple flat charges 232 is transferred to a mandrel 234 having a contour 1610. In Figure 17, the flat charge 232 is shaped to conform to the contour 1610 by the application of pressure and / or heat, resulting in a preform 242. The illustrated complete flat charge 233, having multiple flat charges 232, is shaped to conform to the contour 1610, although in some embodiments not shown, each flat charge 232 is shaped individually to conform to the contour 1610.

[0072] Figure 18 shows a scarf joint 1850 (or other overlap) between spur segments 112 and 114 in an exemplary embodiment. As shown in Figure 18, spur segment 112 includes an end 1834 that overlaps with end 1844 of spur segment 114. The ends overlap along ramps 1832 and 1842 of spur segments 112 and 114. The respective inclination ratios of ramps 1832 and 1842 can be complementary and / or equal so as to maintain a uniform thickness throughout the scarf joint 1850. Ends 1834 and 1844 are shown simply for clarity and do not have flanges 902 and 906. In further embodiments, the scarf joint 1850 is one of several scarf joints that join multiple spur segments together into a single, integrated spur detail. The scarf joint 1850 may be manufactured by co-curing, co-bonding, fastening, or other desired means. Instead of the scarf joint 1850, a stepped lap joint or other suitable joint should be considered.

[0073] Examples In the following embodiments, additional processes, systems, and methods are described from the viewpoint of manufacturing spar details.

[0074] Referring more closely to the drawings, embodiments of the present disclosure can be described in terms of the manufacture and maintenance of an aircraft in Method 1900 shown in Figure 19, and the aircraft 1902 shown in Figure 20. Figure 19 is a flowchart of a method for manufacturing and maintaining an aircraft according to an exemplary embodiment. In the pre-manufacturing stage, Method 1900 may include the specification and design 1904 of the aircraft 1902 and the procurement of materials 1906. In the manufacturing stage, the manufacturing 1908 of the components and subassemblies of the aircraft 1902 and system integration 1910 are carried out. The aircraft 1902 may then be put into operation 1914 after certification and delivery 1912. During its operation by the customer, the aircraft 1902 is scheduled for periodic maintenance and upkeep 1916 (which may also include modifications, reconfigurations, and refurbishments). Apparatus and methods embodied herein may be employed during any suitable stage of one or more of the manufacturing and maintenance described in Method 1900 (e.g., specification and design 1904, procurement of materials 1906, manufacture of components and subassemblies 1908, system integration 1910, authorization and delivery 1912, operation 1914, maintenance and servicing 1916) and / or in 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).

[0075] Each of the processes of Method 1900 may be carried out or performed by a system integrator, a third party, and / or an operator (e.g., a customer). For the purposes of this specification, a system integrator may include, but is not limited to, any number of aircraft manufacturers and major system subcontractors; a third party may include, but is not limited to, any number of vendors, subcontractors, and suppliers; and an operator may be an airline, leasing company, military organization, service organization, etc.

[0076] Figure 20 is a block diagram of an aircraft in an exemplary embodiment. As shown in Figure 20, an aircraft 1902 manufactured by Method 1900 may include a fuselage 1918 with a number of systems 1920 and interior 1922. Examples of systems 1920 include one or more of the propulsion system 1924, electrical system 1926, hydraulic system 1928, and environmental system 1930. Any number of other systems may be included. Although an aerospace embodiment is shown, the principles of the present invention may also be applied to other industries such as the automotive industry.

[0077] As described above, the apparatus and methods embodied herein may be employed in any one or more stages of manufacturing and maintenance described in Method 1900. For example, a component or subassembly corresponding to the manufacturing of components and subassemblies 1908 may be manufactured or produced in a similar manner to a component or subassembly manufactured during the operation of the aircraft 1902 1914. Also, one or more embodiments of apparatus, embodiments of methods, or combinations thereof may be used in the manufacturing of subassemblies 1908 and system integration 1910, for example, by significantly improving the efficiency of the assembly of the aircraft 1902 or by significantly reducing the cost of the aircraft 902. Similarly, one or more embodiments of apparatus, embodiments of methods, or combinations thereof may be used during the operation of the aircraft 1902, for example, during maintenance and upkeep 1916, but not limited to these. Accordingly, the present invention can be used in any stage or any combination thereof as described herein, for example, in specification and design 1904, material procurement 1906, manufacturing of components and subassemblies 1908, system integration 1910, authorization and delivery 1912, operation 1914, maintenance and servicing 1916) and / or in any suitable component of the aircraft 1902 (for example, the airframe 1918, systems 1920, interior 1922, propulsion system 1924, electrical system 1926, hydraulic system 1928, and / or environmental system 1930).

[0078] In one embodiment, a part comprises a portion of the airframe 1918 and is manufactured during the manufacture of components and subassemblies 1908. This part is then incorporated into the aircraft during system integration 1910 and may be used in operation 1914 until it becomes unusable due to wear. Subsequently, during maintenance and servicing 1916, this part may be discarded and replaced with a newly manufactured part. The components and methods of the present invention may be used over the period of manufacture of components and subassemblies 1908 to manufacture new parts.

[0079] Furthermore, this disclosure includes embodiments as described below.

[0080] Clause 1. A method 400 for manufacturing a spar detail 110 of an aircraft 10, comprising: manufacturing a preform 242 of a fiber-reinforced material 301 for spar segments 112, 114, and 116 402; solidifying the preform 242 to form the spar segments 112, 114, and 116 404; and joining the spar segments 112, 114, and 116 together to form a finished spar detail 110.

[0081] Method 400 of Clause 2. Joining together spar segments 112, 114, and 116 406 includes: applying a joining doubler (340) that partially covers the first spar segment 112 and the second spar segment 114; and joining the joining doubler 340 to the first spar segment 112 and the second spar segment 114 406.

[0082] Method 400 of Clause 3, wherein the application of the joining doubler 340 includes placing the joining doubler 340 between the joining region 341 of the first spar segment 112 and the joining region 341 of the second spar segment 114.

[0083] Method 400 of Clause 4, wherein the application of the joining doubler 340 includes sandwiching the joining region 341 of the first spar segment 112 and the joining region 341 of the second spar segment 114 between the front joining doubler and the rear joining doubler.

[0084] Clause 5. Method 400 of Clause 2, 3, or 4, wherein the application of the joining doubler 340 includes positioning the joining doubler 340 so that it extends across the rib intersection 360 of the spur detail 110.

[0085] Clause 6. Applying the joining doubler 340 includes sandwiching a first spar segment 112 and a second spar segment 114 between the joining doubler 340 and the rib 290, in any way of Clause 2-5 400.

[0086] Clause 7. The method 400 of any of Clauses 2-6, wherein joining the joining doubling 340 406 includes solidifying the joining doubling 340 via cobonding in a press clave.

[0087] Clause 8. Joining together spur segments 112, 114, and 116 406 comprises: applying one of an adhesive and / or a resin to spur segments 112, 114, and 116; and forming a joint between spur segments 112, 114, and 116 in the joining region 341 of spur segments 112, 114, and 116, or any method 400 of Clauses 1 to 7.

[0088] Clause 9. A method 400 of any of Clauses 1 to 8, wherein joining spar segments 112, 114, and 116 together 406 comprises forming one of a butt splice, a lap splice, and a scarf splice between the joining region 341 of the first spar segment 112 and the joining region 341 of the second spar segment 114.

[0089] Clause 10. A spar of an aircraft 10 comprising: a first spar segment 112 comprising fiber-reinforced material 301, the first spar segment 112 comprising a bonding region 341; a second spar segment 114 comprising fiber-reinforced material 301-1, the second spar segment 114 comprising a first bonding region 341 and arranged in series with the first spar segment 112; and a bonding doubler 340 covering at least a portion of the bonding region 341 of the first spar segment 112 and the first bonding region 341 of the second spar segment 114, the spar comprising the bonding doubler 340 coupled to the first spar segment 112 and the second spar segment 114.

[0090] Clause 11. The spar of Clause 10, which is either the front spar 110-1 or the rear spar 110-2.

[0091] Clause 12. A spar according to Clause 10 or 11, wherein a joining doubler 340 is sandwiched between the joining region 341 of the first spar segment 112 and the first joining region 341 of the second spar segment 114.

[0092] Clause 13. A spar according to Clause 10, 11, or 12, wherein the joining doubling 340 includes a front joining doubling and a rear joining doubling, with the joining region 341 of the first spar segment 112 and the first joining region 341 of the second spar segment 114 sandwiched between the front joining doubling and the rear joining doubling.

[0093] Clause 14. A spar of any of Clauses 10-13, in which a joining doubling 340 extends across the rib intersection 360 of the spar.

[0094] Clause 15. A spar of any of Clauses 10-14, wherein the joining region 341 of the first spar segment 112 and the first joining region 341 of the second spar segment 114 are sandwiched between a joining doubling 340 and a rib.

[0095] Clause 16. A spar of any of Clauses 10-15, wherein a bonding doubling 340 is co-bonded to the bonding region 341 of the first spar segment 112 and to the first bonding region 341 of the second spar segment 114.

[0096] Clause 17. Any spar of Clauses 10-16, wherein the second spar segment 114 further includes a second bonding region 341-1, and the spar further comprises: a third spar segment 116 containing fiber-reinforced material 301-2, which includes the bonding region 341-1 and is arranged in series with the second spar segment 114; and a second bonding doubler 340-1 covering at least a portion of the bonding region 341-1 of the third spar segment 116 and at least a portion of the second bonding region 341-1 of the second spar segment 114, which is bonded to the second spar segment 114 and the third spar segment 116.

[0097] Article 18. Manufacture of a portion of an aircraft using any of the spars in Articles 10-17.

[0098] Clause 19. An aircraft wing 12 comprising a multi-segment spar detail 110, wherein the spar detail 110 comprises: a first spar segment 112 comprising fiber-reinforced material 301 and including a joint region 341; a second spar segment 114 comprising fiber-reinforced material 301-1 and including a first joint region 341, and arranged in series with the first spar segment 112; and a joint doubler 340 covering at least a portion of the joint region 341 of the first spar segment 112 and the first joint region 341 of the second spar segment 114. An aircraft wing comprising: a joint doubling 340 co-bonded to the second spar segment 112 and the second spar segment 114; a third spar segment 116 comprising fiber-reinforced material 301-2, comprising a joint region 341-1 and arranged in series with the second spar segment 114; and a second joint doubling 340-1 covering at least a portion of the joint region 341-1 of the third spar segment 116 and at least a portion of the second joint region 341-1 of the second spar segment 114, the second joint doubling 340-1 co-bonded to the second spar segment 114 and the third spar segment 116.

[0099] Clause 20. The wing 12 of an aircraft according to Clause 19, wherein at least one of the joining doubling 340 and the second joining doubling 340-1 extends across the rib intersection 360 of the multi-segment spar detail 110.

[0100] Clause 21. A method 500 for manufacturing a spar detail 110 of an aircraft 10, comprising: manufacturing preforms 242 of fiber-reinforced material 301, 301-1 for a first spar segment 112 and a second spar segment 114 502; joining the end 312 of the preform 242 for the first spar segment 112 to the end 321 of the preform 242 for the second spar segment 114 to define a joining area 341 504; applying at least one preform for a joining doubler 340 to the joining area 341 506; and simultaneously solidifying the preforms for the spar segments 112, 114 and the joining doubler 340 to form a portion of the spar detail 110 508.

[0101] Clause 22. Method 500 of Clause 21 for simultaneously solidifying the preform 242 for spur segments 112, 114 and the preform for the bonding doubling 340, comprising: vacuum bagging the preform 242 for spur segments 112, 114 and the preform for the bonding doubling 340 using a vacuum bag; compacting the preform 242 for spur segments 112, 114 and the preform for the bonding doubling 340 through a vacuum bag 252; and heating the preform 242 for spur segments 112, 114 and the preform for the bonding doubling 340.

[0102] The method of Clause 23, applying at least one preform for the joint doubling 340 506, which includes picking up the preform for the joint doubling 340 and positioning it so that the joint doubling 340 extends across the rib intersection 360 of the spar detail 110.

[0103] Method 500 of Clause 24. Applying at least one preform for the joining doubler 340 506 includes laying up the preform for the joining doubler 340 on the joined ends 312, 321 of the first spur segment 112 and the second spur segment 114 such that the joining doubler 340 extends across the rib intersection 360 of the spur detail 110.

[0104] Clause 25. Applying at least one preform for the joint doubling 340 506 includes laying up the preform for the joint doubling 340 across the rib intersection 360 of the spur detail 110 to the opposite side of the first spur segment 112 and the second spur segment 114, in any way of Clauses 21-24 500.

[0105] Clause 26. Simultaneous solidification of preforms 242 for spar segments 112, 114 and bonding doublings 340 508 is to include: arranging dry fiber preforms for the first spar segment 112, the second spar segment 114 and bonding doublings 340; and injecting resin into the dry fibers, any method 500 of Clauses 21-25.

[0106] Clause 27. To define a joint area 341, joining the end 312 of the preform 242 of the first spar segment 112 to the end 321 of the preform 242 of the second spar segment 114 is to be performed by any method 500 of Clauses 21-26, which includes joining the ends 312, 321 using at least one of a lap splice, a butt splice, and a scarf splice.

[0107] Clause 28. Manufacturing a preform 242 502 comprises: distributing a tow of fiber-reinforced material 301 to form a flat charge 232; and shaping the flat charge 232 into a preform 242 having a desired cross-sectional shape, in any method of Clauses 21-27 500.

[0108] Clause 29. Manufacturing a preform 242 502 further comprises: distributing tows of fiber-reinforced material 301 to form flat charges 232; combining multiple flat charges 232 into a complete flat charge 233; and shaping the complete flat charges 233 into a preform 242 having a desired cross-sectional shape, any method 500 of Clauses 21-28.

[0109] Article 30. A portion of an aircraft assembled according to any of Articles 21 to 29.

[0110] Clause 31. A spar detail 110 comprising a preform 242 for a first spar segment 112, wherein the first spar segment 112 includes a joining region 341; a preform 242 for a second spar segment 114, wherein the second spar segment 114 includes a joining region 341; and a preform for a joining doubler 340, wherein the joining regions 341 are arranged in series with respect to each other, and the preforms 242 and the joining doubler 340 are solidified simultaneously while the joining doubler 340 preform covers at least a portion of the joined region 341 to form a portion of the spar detail 110.

[0111] Clause 32. The spar detail 110 of Clause 31, wherein a joining doubling 340 is positioned at the rib intersection 360 of the spar detail 110.

[0112] Clause 33. The spar detail 110 of Clause 31 or 32, wherein the joining doubler 340 is located on the opposite side of the spar segments 112, 114 from the rib intersection 360 of the spar detail 110, and the rib intersection 360 and the joining doubler 340 sandwich a joining region 341.

[0113] Clause 34. A spar detail 110 of Clause 31, 32, or 33, wherein a joined region 341 between preforms 242 defines at least one of a wrap splice, a butt splice, and a scarf splice.

[0114] Clause 35. The preforms of the preform 242 and the bonding doubling 340 include dry fibers, and resin is injected into the dry fibers while the preform for the bonding doubling 340 covers at least a portion of the bonded area 341, as per any of Clauses 31-34, spar detail 110.

[0115] Clause 36. A spar detail 110 according to any of Clauses 31-35, further comprising: a preform 242 for a third spar segment 116, the third spar segment 116 including a joint area 341-1; and a preform for a second joint doubling 340-1 covering at least a portion of the joint area 341-1 of the third spar segment 116 and a portion of the second joint area 341-1 of the second spar segment 114.

[0116] Article 37. Manufacture of a portion of an aircraft 10 using any of the super detailing 110 of Articles 31-36.

[0117] Clause 38. Wing of an aircraft 12: A first spar segment 112 comprising fiber-reinforced material 301, comprising a joint region 341; a second spar segment 114 comprising fiber-reinforced material 301-1, comprising a first joint region 341 and a second joint region 341-1, wherein the first joint region 341 of the second spar segment 114 is arranged in series with the joint region 341 of the first spar segment 112; a joint doubler 340 covering at least a portion of the joint region 341 of the first spar segment 112 and at least a portion of the first joint region 341 of the second spar segment 114, comprising first An aircraft wing 12 comprising: a spar segment 112 and a joint doubling 340 co-cured with the second spar segment 114; a third spar segment 116 comprising a fiber-reinforced material 301-2, the third spar segment 116 comprising a joint region 341-1 arranged in series with the second joint region 341-1 of the second spar segment 114; and a second joint doubling 340-1 covering at least a portion of the joint region 341-1 of the third spar segment 116 and at least a portion of the second joint region 341-1 of the second spar segment 114, the second joint doubling 340-1 co-cured with the second spar segment 114 and the third spar segment 116.

[0118] Clause 39. The wing 12 of an aircraft according to Clause 38, wherein at least one of the joining doubling 340 and the second joining doubling 340-1 extends across the rib intersection 360 of the spar detail 110.

[0119] 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 bonding doubling 340 and the second bonding doubling 340-1 contain dry fibers, and resin is injected into the predetermined positions of the dry fibers prior to co-curing.

[0120] Clause 41. A method 600 for manufacturing a spar detail 110 of an aircraft 10, comprising: manufacturing preforms 242 of fiber-reinforced materials 301, 301-1, 301-2 for spar segments 112, 114, 116 602; solidifying the preforms 242 to form the spar segments 112, 114, 116 604; and applying fasteners 278 to join the spar segments 112, 114, 116 together to form a finished spar detail 110 606.

[0121] Method 600 of Clause 42. Applying fastener 278 606: applying a joining doubler 340 that partially covers the first spur segment 112 and the second spur segment 114; and attaching fastener 278 to the first spur segment 112 and the second spur segment 114 through the joining doubler 340.

[0122] Clause 43. Method 600 of Clause 42, wherein the joining doubler 340 is applied by sandwiching a first spar segment 112 and a second spar segment 114 between the joining doublers 340.

[0123] Method 600 of Clause 42, wherein the application of the joining doubler 340 includes positioning the joining doubler 340 so that it extends across the rib intersection 360 of the spur detail 110.

[0124] Method 600 of Clause 42, wherein applying the joining doubler 340 includes positioning the joining doubler 340 so that it extends from the rib intersection 360 of the spur detail 110 to the opposite side of the spur detail 110.

[0125] Method 600 of Clause 46, further comprising applying a second joining doubler 340-1 that partially covers the second spur segment 114 and the third spur segment 116; and attaching a fastener 278 to the second spur segment 114 and the third spur segment 116 through the second joining doubler 340-1.

[0126] Clause 47. Manufacturing a preform 242 602 comprises: distributing a tow of fiber-reinforced material 301 to form a flat charge 232; and shaping the flat charge 232 into a preform 242 having a desired cross-sectional shape, any method 600 of Clauses 41-46.

[0127] Clause 48. Manufacturing a preform 242 602 includes: distributing tows of fiber-reinforced material 301 to form flat charges 232; combining multiple flat charges 232 into a complete flat charge 233; and shaping the complete flat charges 233 into a preform 242 having a desired cross-sectional shape, any method 600 of Clauses 41-47.

[0128] Article 49. A portion of an aircraft 10 assembled by any of the methods 600 of Articles 41-48.

[0129] Clause 50. Spur detail 110 of an aircraft 10, comprising: a first spur segment 112 comprising fiber-reinforced material 301; a second spur segment 114 comprising fiber-reinforced material 301-1 and arranged in series with the first spur segment 112; a joining doubler 340 covering a joining region 341 between the first spur segment 112 and the second spur segment 114; and a fastener 278 attached through the joining doubler 340, the first spur segment 112, and the second spur segment 114 to form at least a portion of the spur detail 110.

[0130] Clause 51. Spar detail 110 of Clause 50, wherein a joining doubling 340 is positioned across the rib intersection 360 and between the first spar segment 112 and the second spar segment 114.

[0131] Clause 52. Spur detail 110 of Clause 50 or 51, wherein a joining doubling 340 is located on the opposite side of the first spur segment 112 and the second spur segment 114 from the rib intersection 360.

[0132] Clause 53. A spar detail 110 of Clause 50, 51, or 52, comprising: a third spar segment 116 comprising fiber-reinforced material 301-2 and arranged in series with a second spar segment 114; a second joining doubler 340-1 covering a joining region 341-1 between the second spar segment 114 and the third spar segment 116; and a fastener 278 attached through the second joining doubler 340-1, the second spar segment 114, and the third spar segment 116 to form at least a portion of the spar detail 110.

[0133] Clause 54. A spur detail 110 of any of Clauses 50-53, in which a first spur segment 112 and a second spur segment 114 are sandwiched between a joining doubler 340.

[0134] Article 55. Manufacture of a portion of an aircraft using the super detail 110 of Article 50.

[0135] Clause 56. An aircraft wing 12 comprising: a first spar segment 112 comprising fiber-reinforced material 301, the first spar segment 112 comprising a joint region 341; a second spar segment 114 comprising fiber-reinforced material 301, comprising a first joint region 341 and a second joint region 341-1, wherein the first joint region 341 of the second spar segment 114 is arranged in series with the joint region 341 of the first spar segment 112; a joint doubler 340 covering at least a portion of the joint region 341 of the first spar segment 112 and at least a portion of the first joint region 341 of the second spar segment 114; and a fastener 278 attached through the joint doubler 340, the joint region 341 of the first spar segment 112, and the first joint region 341 of the second spar segment 114 to form at least a portion of the spar detail 110.

[0136] Clause 57. An aircraft wing 12 according to Clause 56, comprising: a third spar segment 116 containing fiber-reinforced material 301, the third spar segment 116 including a joint region 341-1 arranged in series with a second joint region 341-1 of a second spar segment 114; a second joint doubler 340-1 covering at least a portion of the joint region 341-1 of the third spar segment 116 and at least a portion of the second joint region 341-1 of the second spar segment 114; and a fastener 278 attached through the second joint doubler 340-1, the second joint region 341-1 of the second spar segment 114, and the joint region 341-1 of the third spar segment 116 to form at least a portion of the spar detail 110.

[0137] Article 58. An aircraft wing 12 according to Article 57, wherein at least one of the joining doubling 340 and the second joining doubling 340-1 extends across the rib intersection 360.

[0138] Clause 59. The wing of an aircraft according to Clause 57 or 58, wherein a joining doubler 340 is located opposite the first spar segment 112 and the second spar segment 114 from the rib intersection 360; and at least one of the second joining doubler 340-1 is located opposite the second spar segment 114 and the third spar segment 116 from the rib intersection 360.

[0139] Article 60. Manufacture of a portion of an aircraft 10 using the wings 12 of any of the aircraft specified in Articles 57-59.

[0140] Clause 61. A method 700 for manufacturing a spar of an aircraft 10, comprising: manufacturing a preform 242 of fiber-reinforced material 301 for spar segments 112, 114, 116 702, wherein at least one of the spar segments 112, 114, 116 includes a kink 370, and each kink 370 is completely contained within the preform 242; solidifying the preform 242 to form the spar segments 112, 114, 116 704; and assembling the spar segments 112, 114, 116 together to form a finished spar detail 110 exhibiting at least one of the kinks 370 706.

[0141] Clause 62. The method of Clause 61, wherein each kink 370 includes a bending point 381 where the intersection of the first neutral axis 350 and the second neutral axis 350-1 of the spur is located.

[0142] Clause 63. Method 700 of Clause 61 or 62, wherein manufacturing the preform 242 702 includes changing the axial direction of the preform 242 at each kink 370.

[0143] Method 700 of Clause 64. Manufacturing the preform 242 702 comprises changing the axial direction of the preform 242 at each kink 370, wherein the axial change is a bending angle of 2 to 10 degrees.

[0144] Clause 65. Each preform 242 includes a first end and a second end opposite to the first end, and manufacturing the preform 242 702 includes positioning each kink 370 at least about 30 cm (1 foot) from the end of the preform 242, in any way of Clauses 61-64 700.

[0145] Clause 66. Each preform 242 includes at least one bonding region 341, 341-1, and assembling the spar segments 112, 114, 116 706 is done in any way of Clauses 61-65 700, which includes applying bonding doublings 340, 340-1 to the spar segments 112, 114, 116 within the bonding region 341, 341-1.

[0146] Method of Clause 67. Applying splice doublings 340, 340-1 to spar segments 112, 114, 116: forming at least one of lap splices, butt splices and scarf splices between adjacent splice regions 341, 341-1; and attaching splice doublings 341, 341-1 to splice regions 340, 341-1 using one of co-curing, co-bonding and fastener 278, as per the method of Clause 66 700.

[0147] Method 700 of Clause 68. Applying joining doublings 340, 340-1 to spar segments 112, 114, 116 within joining regions 341, 341-1, including positioning joining doublings 340, 340-1 across the rib intersections 360 of the spar.

[0148] Method 700 of Clause 69. Applying joining doublings 340, 340-1 to spar segments 112, 114, 116 within joining regions 341, 341-1, including sandwiching joining regions 341, 341-1 between joining doublings 340, 340-1 and ribs.

[0149] Clause 70. Applying joining doublers 340, 340-1 to spar segments 112, 114, 116 within joining regions 341, 341-1, including sandwiching joining regions 341-341-1 between a front joining doubler and a rear joining doubler, according to any method of Clauses 66-69 700.

[0150] Clause 71. Manufacturing a preform 242 of fiber-reinforced material 301 for spar segments 112, 114, 116, each preform 242 comprising at least one bonding region 341, 341-1, and comprising placing a kink 370 outside the bonding region 341-, 341-1 of the spar segments 112, 114, 116, in any method of Clauses 61-70 700.

[0151] Clause 72. Aircraft spar detail 110 comprising: a first spar segment 112 comprising fiber-reinforced material 301 and at least one joint region 341; a second spar segment 114 comprising fiber-reinforced material 301, at least one joint region 341, and a kink 370 outside the joint region 341, wherein the respective joint regions 341 are arranged in series with respect to each other; and a joint doubler 340 covering at least a portion of the joint region 341 of the first spar segment 112 and at least a portion of the corresponding joint region 341 of the second spar segment 114, the joint doubler 340 being connected to the first spar segment 112 and the second spar segment 114.

[0152] Clause 73. The spar detail 110 of the aircraft in Clause 72 includes a kink 370 which includes a variation in the axial direction of the spar detail 110.

[0153] Clause 74. Aircraft spar detail 110 of Clause 73, where the axial change is 2 to 10 degrees.

[0154] Clause 75. The spar detail 110 of an aircraft according to Clause 72, 73, or 74, is separated from the end of the second spar segment 114 by more than 30 cm (1 foot).

[0155] Clause 76. Spar detail 110 of an aircraft according to any of Clauses 72-75, wherein the first spar segment 112 includes a kink 370 outside the joint area 341.

[0156] Clause 77. Spar detail 110 of any aircraft according to Clauses 72-76, in which a joining doubling 340 is located across a rib intersection 360.

[0157] Clause 78. The spar detail of the aircraft according to Clause 77, wherein the joining doubling 340 includes: a front joining doubling and a rear joining doubling that sandwich the joining region 341; and one of the joining doublings 340 positioned between the joining doubling 340 and the rib to sandwich the joining region 341.

[0158] Clause 79. A spar detail 110 of an aircraft according to any of Clauses 72-78, comprising: a third spar segment 116 including fiber-reinforced material 301, at least one joint region 341-1, and a kink 370-1 outside the joint region 341, wherein the joint region 341-1 is arranged in series with the corresponding joint region 341-1 of a second spar segment 114; and a second joint doubling 340-1 covering at least a portion of the joint region 341-1 of the third spar segment 116 and at least a portion of the corresponding joint region 341-1 of the second spar segment 114, the second joint doubling 340-1 being connected to the second spar segment 114 and the third spar segment 116.

[0159] Clause 80. Wing of an aircraft 12, comprising: a first spar segment 112 comprising fiber-reinforced material 301, the first spar segment 112 comprising a joint region 341; a second spar segment 114 comprising fiber-reinforced material 301, comprising a first joint region 341 and a second joint region 341-1, wherein the first joint region 341 of the second spar segment 114 is arranged in series with the joint region 341 of the first spar segment 112; a joint doubler 340 covering at least a portion of the joint region 341 of the first spar segment 112 and at least a portion of the first joint region 341 of the second spar segment 114; a third spar segment 116 comprising fiber-reinforced material 301, An aircraft wing 12 comprising: a third spar segment 116 including a joining region 341-1 positioned in series with the second joining region 341-1 of two spar segments 114; a second joining doubler 340-1 covering at least a portion of the joining region 341-1 of the third spar segment 116 and at least a portion of the second joining region 341-1 of the second spar segment 114; and at least one kink 370 in one or more of the first spar segment 112, the second spar segment 114, and the third spar segment 116, which is located outside the joining regions 341, 341-1 and includes at least one kink 370 that includes axial variation in the spar segments 112, 114, 116 in which the kink 370 is located.

[0160] Clause 81. A method 800 for manufacturing a spar detail 110 of an aircraft 10, comprising: manufacturing a preform 242 for a first spar segment 112 802, wherein the preform 242 includes a sub-kink 371-1 adjacent to one end 112 of the first spar segment 802; manufacturing a preform 242 for a second spar segment 114 802, wherein the preform 242 includes a sub-kink 371 adjacent to one end 114 of the second spar segment 802; aligning the ends of the preform 242 such that the sub-kinks 371-1, 371 are adjacent to each other within a joining area 341 804; and joining the spar segments 112, 114 together in the joining area 341 to form at least a portion of the spar detail 110 exhibiting a kink 370 806.

[0161] Method 800 of Clause 82, 806, of joining spar segments 112, 114 in a joining region 341 using at least one of co-curing, co-bonding, attachment of a joining doubling 340, and attachment of a fastener 278.

[0162] Method 800 of Clause 83. Manufacturing a preform 242 802 comprises manufacturing the preform 242 such that each sub-kink 371-1, 371 has an equal angular deviation.

[0163] Method 800 of Clause 84. Manufacturing a preform 242 802 includes manufacturing the preform 242 such that the sub-kinks 371-1, 371 have angular deviations that are not equal.

[0164] Article 85. Manufacturing a preform 242 802 includes any method of Articles 81-84, wherein the preform 242 is manufactured such that the sub-kinks 371-1, 371 collectively change the axial direction of the spur detail 110 by 2 to 10 degrees.

[0165] Clause 86. Aligning the ends of the preform 242 804 includes forming one of a lap splice, a butt splice, and a scarf splice in the joining area 341 with the ends of the preform 242, method 800 of any of Clauses 81-85.

[0166] Clause 87. Joining together spar segments 112, 114 806 includes applying a joining doubler 340 to the joining region 341 opposite the rib intersection 360, in any method 800 of Clauses 81-86.

[0167] Clause 88. Joining together spar segments 112, 114 806 includes inserting sub-kinks 371-1, 371 between the joining doubling 340 and the rib, in any method of Clauses 81-87 800.

[0168] Clause 89. Joining spar segments 112, 114 together 806 includes inserting sub-kinks 371-1, 371 between the front joining doubling and the rear joining doubling, in any method 800 of Clauses 81-88.

[0169] Article 90. A portion of an aircraft 10 assembled by any of the methods of Articles 81-89.

[0170] Clause 91. Spar detail 110 of an aircraft 10, comprising: a first spar segment 112 comprising fiber-reinforced material 301 and having a sub-kink 371-1 located at one end; a second spar segment 114 comprising fiber-reinforced material 301 and having a sub-kink 371 located at one end, wherein the end of the first spar segment 112 having the sub-kink 371-1 is 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 end defines a joining region 341; and a joining doubling 340 for structurally joining the first spar segment 112 and the second spar segment 114 within the joining region 341.

[0171] Spur detail 110 of Clause 92. A sub-kink 371-1 associated with the first spur segment 112 is located within approximately 30 cm (1 foot) of the end of the first spur segment 112; and a sub-kink 371 associated with the second spur segment 114 is located within approximately 30 cm (1 foot) of the end of the second spur segment 114.

[0172] Clause 93. Each sub-kink 371-1, 371 changes the axial direction of the spar detail 110 by half the amount of kink 370, according to Clause 91 or 92.

[0173] Clause 94. Sub-kinks 371-1 and 371 together cause a change in the axial direction of the spur detail 110 by 2 to 10 degrees, as per Clause 91, 92, or 93.

[0174] Clause 95. A spar detail 110 of any of Clauses 91-94, wherein a joining doubling 340 is located on the opposite side of a rib intersection 360 defined in the spar detail 110.

[0175] Clause 96. A spar detail 110 of any of Clauses 91-95, wherein a joining doubling 340 is positioned to sandwich the joining area 341 and the sub-kinks 371-1, 371.

[0176] Clause 97. To structurally join the first spur segment 112 and the second spur segment 114, the joining doubling 340 is one of the following: co-bonded to the first spur segment 112 and the second spur segment 114, co-cured, and attached by fastener 278, in any spur detail 110 of Clauses 91-96.

[0177] Article 98. Manufacture of a portion of an aircraft 10 using any of the super detailing 110s in Articles 91-97.

[0178] Clause 99. Wing of an aircraft 12, comprising: a first spar segment 112 comprising fiber-reinforced material 301 and having a sub-kink 371-1 located at one end; a second spar segment 114 comprising fiber-reinforced material 301 and having a sub-kink 371 located at one end, wherein the end of the first spar segment 112 having the sub-kink 371-1 is 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 end defines a joint region 341; An aircraft wing 12 comprising: a joining doubling 340 for structurally joining a first spar segment 112 and a second spar segment 114 within a joining region 341; a third spar segment 116 comprising fiber-reinforced material 301, the third spar segment 116 including a joining region 341-1 positioned in series with the second joining region 341-1 at the opposite end of the second spar segment 114; and a second joining doubling 340-1 covering at least a portion of the joining region 341-1 of the third spar segment 116 and at least a portion of the second joining region 341-1 of the second spar segment 114.

[0179] Clause 100. The joining doubling 340 of the aircraft wing 12 of Clause 99 is one of the following: co-bonded to a first spar segment 112 and a second spar segment 114, co-cured, and attached by fastener 278.

[0180] Any of the various control elements (e.g., electrical or electronic components) shown in the illustrations or described herein may be implemented as hardware, processor-implemented software, processor-implemented firmware, or any combination thereof. For example, one element may be implemented as dedicated hardware. Dedicated hardware elements may be referred to as “processors,” “controllers,” or any similar terms. Where provided by processors, functions may be provided by a single dedicated processor, a single shared processor, or by multiple individual processors, some of which may be shared. Furthermore, the explicit use of the terms “processors” or “controllers” should not be interpreted as referring only to hardware capable of running software, but implicitly includes, but is not limited to, digital signal processor (DSP) hardware, network processors, application-specific integrated circuits (ASICs) or other circuits, field-programmable gate arrays (FPGAs), read-only memory (ROM) for software storage, random-access memory (RAM), non-volatile memory, logic, or any other physical hardware components or modules.

[0181] Furthermore, control elements may be implemented as instructions that can be executed by a processor or computer to perform the function of the element. Some examples of instructions are software, program code, and firmware. When executed by a processor, an instruction is operable to instruct the processor to perform the function of the element. Instructions may be stored in a storage device that can be read by the processor. Some examples of storage devices are digital or solid-state memory, magnetic storage media such as magnetic disks and magnetic tapes, hard drives, or optically readable digital data storage media.

[0182] While specific embodiments are described herein, the scope of this disclosure is not limited to those specific embodiments. The scope of this disclosure is defined by the claims.

[0183] This disclosure further includes the following provisions:

[0184] Clause Clause 1. A method for manufacturing spar details for an aircraft, which: To manufacture preforms of fiber-reinforced material for spar segments; Solidifying the preform to form spar segments; and To form a completed spar detail, spar segments are joined together. A method that includes this.

[0185] Clause 2. Spar segments can be joined together: Applying a joining doubler that partially covers the first spar segment and the second spar segment; and Joining a junction doubler to the first spar segment and the second spar segment. The method of Clause 1, including the method of Clause 1.

[0186] Clause 3. The method of Clause 2, wherein the application of a joining doubler includes placing a joining doubler between the joining region of the first spar segment and the joining region of the second spar segment.

[0187] Clause 4. The method of Clause 2, wherein the application of the joining doubler includes sandwiching the joining region of the first spar segment and the joining region of the second spar segment between the front joining doubler and the rear joining doubler.

[0188] Clause 5. The method of Clause 2, wherein the application of a joining doubler includes positioning the joining doubler so that it extends across the rib intersection of the spar detail.

[0189] Clause 6. The method of Clause 2, wherein the application of the joining doubler includes inserting a first spar segment and a second spar segment between the joining doubler and the rib.

[0190] Clause 7. The method of Clause 2, wherein joining the joining doublings includes solidifying the joining doublings via cobonding in a preclave.

[0191] Clause 8. Spar segments can be joined together: Applying either an adhesive or a resin to the spar segment; and Forming a joint between spar segments within the joining region of a spar segment. The method of Clause 1, including the method of Clause 1.

[0192] Clause 9. The method of Clause 1, wherein joining spar segments together comprises forming one of a butt splice, a wrap splice, and a scarf splice between the joining region of a first spar segment and the joining region of a second spar segment.

[0193] Article 10. Aircraft spar: A first spar segment comprising a fiber-reinforced material, the first spar segment comprising a bonding region; A second spar segment comprising a fiber-reinforced material, the second spar segment comprising a first bonding region and arranged in series with the first spar segment; and A joining doubler covering at least a portion of the joining region of a first spar segment and the first joining region of a second spar segment, wherein the joining doubler is coupled to the first spar segment and the second spar segment. A spar equipped with a spar.

[0194] Clause 11. The spar of Clause 10, which is either the front spar 110-1 or the rear spar 110-2.

[0195] Clause 12. A spar according to Clause 10, wherein a joining doubler is sandwiched between the joining region of a first spar segment and the first joining region of a second spar segment.

[0196] Clause 13. The spar of Clause 10, wherein the joining doubler includes a front joining doubler and a rear joining doubler, with the joining region of a first spar segment and the first joining region of a second spar segment sandwiched between the front joining doubler and the rear joining doubler.

[0197] Clause 14. A spar of Clause 10, in which a joining doubler extends across the rib intersection of the spar.

[0198] Clause 15. A spar according to Clause 10, wherein the joining region of the first spar segment and the first joining region of the second spar segment are sandwiched between a joining doubling and a rib.

[0199] Clause 16. A spar according to Clause 10, wherein a bonding doubling is co-bonded to the bonding region of a first spar segment and to the first bonding region of a second spar segment.

[0200] Clause 17. The second spar segment further includes a second joining region, and the spar is: A third spar segment comprising a fiber-reinforced material, the third spar segment comprising a bonding region and arranged in series with a second spar segment; and A second joining doubler covering at least a portion of the joining region of a third spar segment and at least a portion of the second joining region of a second spar segment, wherein the second joining doubler is coupled to the second spar segment and the third spar segment. A spar of clause 10 further includes the following.

[0201] Article 18. Manufacture of an aircraft part using the spars of Article 10.

[0202] Clause 19. Wings of aircraft having multi-segment spar details, wherein the spar details are: A first spar segment comprising a fiber-reinforced material, the first spar segment comprising a bonding region; A second spar segment comprising a fiber-reinforced material, comprising a first bonding region, and arranged in series with the first spar segment; A junction doubler covering at least a portion of the junction region of a first spar segment and the first junction region of a second spar segment, wherein the junction doubler is co-bonded to the first spar segment and the second spar segment; A third spar segment comprising a fiber-reinforced material, the third spar segment including a bonding region and arranged in series with a second spar segment; and A second bonding doubling covering at least a portion of the bonding region of a third spar segment and at least a portion of the second bonding region of a second spar segment, wherein the second bonding doubling is co-bonded to the second spar segment and the third spar segment. The wings of an aircraft, including the wings themselves.

[0203] Clause 20. The wing of an aircraft according to Clause 19, wherein at least one of the joining doubling and the second joining doubling extends across the rib intersection of a plurality of segmented spar details.

[0204] co-cure Article 21. A method for manufacturing spar details of an aircraft: To manufacture preforms of fiber-reinforced material for a first spar segment and a second spar segment; Joining the end of the first spar segment preform to the end of the second spar segment preform in order to define the joining area; Applying at least one preform for bonding doublers to the bonding region; and To form a part of the spar detail, the spar segment and the preform for the joining doubler are solidified simultaneously. A method that includes this.

[0205] Clause 22. The preforms for spar segments and preforms for bonding doublings may be solidified simultaneously: Vacuum bagging of preforms for spar segments and preforms for bonding doublings using vacuum bags; Compacting preforms for spar segments and preforms for bonding doublings via vacuum bags; and Heating the preforms for spar segments and the preforms for bonding doublings. The method of Article 21, including the method of Article 21.

[0206] Article 23. The method of Clause 21, wherein applying at least one preform for a joining doubler includes picking up the preform for the joining doubler and positioning it so that the joining doubler extends across the rib intersection of the spar detail.

[0207] Article 24. The method of Clause 21, wherein applying at least one preform for a joining doubler includes laying up the preform for a joining doubler on the joined ends of a first spar segment and a second spar segment such that the joining doubler extends across the rib intersection of the spar detail.

[0208] Article 25. The method of Clause 21, wherein applying at least one preform for a joint doubler includes laying up the preform for the joint doubler across the rib intersection of the spar detail from the first spar segment to the opposite side of the second spar segment.

[0209] Clause 26. The preforms for spar segments and joint doublings may be solidified simultaneously: Arranging the first spar segment, the second spar segment, and the dry fiber preform for the bonding doubling; and Injecting resin into dry fibers The method of Article 21, including the method of Article 21.

[0210] Clause 27. The method of Clause 21, wherein joining the end of the preform of a first spar segment to the end of the preform of a second spar segment in order to define a joining area includes joining the ends using at least one of a lap splice, a butt splice, and a scarf splice.

[0211] Clause 28. To manufacture preforms: Distributing tows of fiber-reinforced material to form flat charges; and Shaping a flat charge into a preform having a desired cross-sectional shape. The method of Article 21, including the method of Article 21.

[0212] Clause 29. To manufacture preforms: Distributing the tow of fiber-reinforced material to form a flat charge; Combining multiple flat charges into a complete flat charge, and Shaping a complete flat charge into a preform having the desired cross-sectional shape. The method of Article 21, including the method of Article 21.

[0213] Article 30. A portion of an aircraft assembled in accordance with the method of Article 21.

[0214] Article 31. A preform for a first spar segment, wherein the first spar segment includes a bonding region; A preform for a second spar segment, wherein the second spar segment includes a bonding region; and Joining Doubling Preform Equipped with, The junction regions are arranged in series with respect to each other. The preform and the bonding doubling preform are cured simultaneously while the bonding doubling preform covers at least a portion of the bonded area in order to form a part of the spar detail. Super detail.

[0215] Clause 32. A spar detail according to Clause 31, wherein a joining doubler is located at the rib intersection of the spar detail.

[0216] Clause 33. A spar detail according to Clause 31, wherein the joining doubler is located on the side of the spar segment opposite to the rib intersection of the spar detail, and the rib intersection and the joining doubler straddle the joining region.

[0217] Clause 34. A spar detail of Clause 31 in which a joined area between preforms defines at least one of a wrap splice, a butt splice, and a scarf splice.

[0218] Clause 35. The spar detail of Clause 31, wherein the preform and the bonding doubling preform include dry fibers, and resin is injected into the dry fibers while the bonding doubling preform covers at least a portion of the bonded area (341).

[0219] Article 36. A preform (242) for a third spar segment (330), wherein the third spar segment includes a bonding region; and A second bonding doubling preform (242) covering at least a portion of the bonding region (341) of the third spar segment and a portion of the second bonding region of the second spar segment. Further details of Clause 31 include:

[0220] Article 37. Manufacturing of a portion of an aircraft using the spar detailing of Article 31.

[0221] Article 38. Wings of an aircraft: A first spar segment comprising a fiber-reinforced material, the first spar segment comprising a bonding region; A second spar segment comprising a fiber-reinforced material, comprising a first bonding region and a second bonding region, wherein the first bonding region of the second spar segment is arranged in series with the bonding region of the first spar segment; A bonding doubling covering at least a portion of the bonding region of a first spar segment and at least a portion of the first bonding region of a second spar segment, the bonding doubling being co-cured with the first spar segment and the second spar segment; A third spar segment comprising a fiber-reinforced material, the third spar segment comprising a bonding region arranged in series with a second bonding region of a second spar segment; and A second bonding doubling covering at least a portion of the bonding region of a third spar segment and at least a portion of the second bonding region of a second spar segment, wherein the second bonding doubling is co-cured with the second spar segment and the third spar segment. The wing of an aircraft equipped with [a specific feature / feature].

[0222] Clause 39. The wing of an aircraft according to Clause 38, wherein at least one of the joining doubling and the second joining doubling extends across the rib intersection of the spar detail.

[0223] Clause 40. The wing of an aircraft according to Clause 38, wherein, prior to co-curing, the first spar segment, the second spar segment, the third spar segment, the bonding doubling and the second bonding doubling contain dry fibers, and resin is injected into the predetermined positions of the dry fibers prior to co-curing.

[0224] Fastener Article 41. A method for manufacturing spar details of an aircraft: To manufacture preforms of fiber-reinforced material for spar segments; Solidifying the preform to form spar segments; and Applying fasteners to connect spar segments together to form a completed spar detail. A method that includes this.

[0225] Clause 42. The fastener may be applied: Applying a joining doubler that partially covers the first spar segment and the second spar segment; and Attaching fasteners through connecting doublers in the first and second spar segments. The method of Article 41, including the method of Article 41.

[0226] Clause 43. The method of Clause 42, wherein the application of the joining doubler includes sandwiching a first spar segment and a second spar segment between the joining doublers.

[0227] Clause 44. The method of Clause 42, wherein the application of a joining doubler includes positioning the joining doubler so that it extends across the rib intersection of the spar detail.

[0228] Clause 45. The method of Clause 42, wherein the application of a joining doubler includes positioning the joining doubler so that it extends from the rib intersection of the spar detail to the opposite side of the spar detail.

[0229] Article 46. Applying a second joining doubler that partially covers the second and third spar segments; and Attaching fasteners through the second joining doubler in the second spar segment and the third spar segment. The methods of Article 42, further including the following.

[0230] Clause 47. To manufacture preforms: Distributing tows of fiber-reinforced material to form flat charges; and Shaping a flat charge into a preform having a desired cross-sectional shape. The method of Article 41, including the method of Article 41.

[0231] Clause 48. To manufacture preforms: Distributing the tow of fiber-reinforced material to form a flat charge; Combining multiple flat charges into a complete flat charge, and Shaping a complete flat charge into a preform having the desired cross-sectional shape. The method of Article 21, including the method of Article 21.

[0232] Article 49. A part of an aircraft assembled in any of the manner described in Articles 41 to 48.

[0233] Clause 50. Aircraft spar details: A first spar segment containing fiber-reinforced material; A second spar segment comprising fiber-reinforced material and arranged in series with a first spar segment; A joining doubler covering the joining region between the first spar segment and the second spar segment; and A fastener attached through a joint double, a first spar segment, and a second spar segment to form at least a portion of the spar detail. Super detailing with added features.

[0234] Clause 51. Spar detail of Clause 50, wherein a joining doubler is positioned across the rib intersection and between the first spar segment and the second spar segment.

[0235] Clause 52. Spar detail of Clause 50, wherein a joining doubler is located on the opposite side of the first and second spar segments from the rib intersection.

[0236] Article 53. A third spar segment containing fiber-reinforced material and arranged in series with the second spar segment; A second joining doubler covering the joining region between the second spar segment and the third spar segment; and A second joining doubling, a second spar segment, and a fastener attached through the third spar segment to form at least a portion of the spar Further features include the spar details of Clause 50.

[0237] Clause 54. Spur detail of Clause 50, wherein a first spur segment and a second spur segment are sandwiched between joining doublers.

[0238] Clause 55. Manufacture of a part of an aircraft using any of the spar details in Clauses 50-54.

[0239] Article 56. Wings of an aircraft: A first spar segment comprising a fiber-reinforced material, the first spar segment comprising a bonding region; A second spar segment comprising a fiber-reinforced material, comprising a first bonding region and a second bonding region, wherein the first bonding region of the second spar segment is arranged in series with the bonding region of the first spar segment; A joining doubler covering at least a portion of the joining region of the first spar segment and at least a portion of the first joining region of the second spar segment; and A fastener is attached through a joining double, the joining region of the first spar segment, and the first joining region of the second spar segment to form at least a portion of the spar detail. The wing of an aircraft equipped with [a specific feature / feature].

[0240] Article 57. A third spar segment comprising a fiber-reinforced material, the third spar segment comprising a bonding region arranged in series with a second bonding region of a second spar segment; A second joining doubling covering at least a portion of the joining region of the third spar segment and at least a portion of the second joining region of the second spar segment; and A fastener attached through a second splice doubler, a second joint area of the second spar segment, and a joint area of the third spar segment to form at least a portion of the spar detail The aircraft wing of clause 56, further comprising the same

[0241] The aircraft wing of clause 57, wherein at least one of the splice doubler and the second splice doubler extends across the rib intersection

[0242] Clause 59 The splice doubler is disposed on the opposite side of the first spar segment and the second spar segment from the rib intersection; and The second splice doubler is disposed on the opposite side of the second spar segment and the third spar segment from the rib intersection The aircraft wing of clause 57, which is at least one of the above

[0243] Clause 60. Manufacture of a part of an aircraft using the aircraft wing of any one of clauses 57 - 59

[0244] Spark kink Clause 61. A method for manufacturing a spar of an aircraft, comprising: Manufacturing a preform of a fiber - reinforced material for a spar segment, wherein at least one of the spar segments includes a kink and each kink is entirely contained within the preform; Solidifying the preform to form the spar segment; and Assembling the spar segments together to form a finished spar detail presenting at least one of the kinks The method including the above

[0245] Clause 62. The method of clause 61, wherein each kink includes a bending point where an intersection of a first neutral axis and a second neutral axis of the spar exists

[0246] Clause 63. The method of Clause 61, wherein manufacturing the preform includes changing the axial direction of the preform at each kink.

[0247] Clause 64. The method of Clause 61, wherein manufacturing the preform includes changing the axial direction of the preform at each kink, and the axial change is a bending angle of 2 to 10 degrees.

[0248] Clause 65. Each preform includes a first end and a second end opposite the first end, and the method of Clause 61, wherein manufacturing the preform includes disposing each kink at least 30 cm (1 foot) from the ends of the preform.

[0249] Clause 66. Each preform includes at least one joining region, and the method of Clause 61, wherein assembling the sparse segments includes applying joining doublers to the sparse segments within the joining region.

[0250] Clause 67. Applying joining doublers to the sparse segments includes: forming at least one of a lap splice, a butt splice, and a scarf splice between adjacent joining regions; and attaching the joining doublers to the joining regions using one of co-curing, co-bonding, and fasteners The method of Clause 66.

[0251] Clause 68. The method of Clause 66, wherein applying joining doublers to the sparse segments within the joining region includes disposing the joining doublers across the rib intersections of the spars.

[0252] Clause 69. The method of Clause 66, wherein applying joining doublers to the sparse segments within the joining region includes sandwiching the joining region between the joining doublers and the ribs.

[0253] Clause 70. The method of Clause 66, wherein applying joining doublers to the sparse segments within the joining region includes sandwiching the joining region between a front joining doubler and a rear joining doubler.

[0254] Clause 71. The method of Clause 61, wherein each preform includes at least one bonding region, and the production of a fiber-reinforced material preform for a spar segment includes locating a kink outside the bonding region of the spar segment.

[0255] Clause 72. Aircraft spar details: A first spar segment comprising a fiber-reinforced material and at least one bonding region; A second spar segment comprising a fiber-reinforced material, at least one bonding region, and a kink outside the bonding region, wherein each bonding region is arranged in series with respect to the others; and A joining doubler covering at least a portion of the joining region of a first spar segment and at least a portion of the corresponding joining region of a second spar segment, the joining doubler being connected to the first spar segment and the second spar segment. The aircraft's super detail, featuring...

[0256] Clause 73. A kink in the spar detail of an aircraft, including a variation in the axial direction of the spar detail, as per Clause 72.

[0257] Clause 74. Spar detail of an aircraft according to Clause 73, where the axial change is between 2 and 10 degrees.

[0258] Clause 75. Aircraft spar detail of Clause 72, where the kink is separated from the end of the second spar segment by more than 30 cm (1 foot).

[0259] Clause 76. Spar detail of an aircraft according to Clause 72, wherein the first spar segment includes a kink outside the joint area.

[0260] Clause 77. Spar detail of an aircraft according to Clause 72, where a joining doubling is located across a rib intersection.

[0261] Article 78. Joining doublers: A front joint doubler and a rear joint doubler sandwiching a joint area; and A joint doubler arranged so as to sandwich a joint area between the joint doubler and a rib One of which is included in the aircraft spar detail of clause 77.

[0262] Clause 79. A third spar segment including a fiber reinforced material, at least one joint area, and a kink outside the joint area, the joint area being arranged in series with a corresponding joint area of a second spar segment; and A second joint doubler covering at least a part of the joint area of the third spar segment and at least a part of the corresponding joint area of the second spar segment, the second joint doubler being connected to the second spar segment and the third spar segment Further comprising the aircraft spar detail of clause 72.

[0263] Clause 80. An aircraft wing comprising: A first spar segment including a fiber reinforced material, the first spar segment including a joint area; A second spar segment including a fiber reinforced material, including a first joint area and a second joint area, the first joint area of the second spar segment being arranged in series with the joint area of the first spar segment, the second spar segment; A joint doubler covering at least a part of the joint area of the first spar segment and at least a part of the first joint area of the second spar segment; A third spar segment including a fiber reinforced material, including a joint area arranged in series with the second joint area of the second spar segment, the third spar segment; A second joint doubler covering at least a part of the joint area of the third spar segment and at least a part of the second joint area of the second spar segment; and At least one kink of one or more of the first spar segment, the second spar segment, and the third spar segment, which is located outside the joint region and includes a change in the axial direction of the spar segment in which the kink is located. The wings of an aircraft equipped with [a specific feature / feature].

[0264] Split lamp Article 81. A method for manufacturing spar details of an aircraft, which: Manufacturing a preform for a first spar segment, wherein the preform includes a sub-kink adjacent to one end of the first spar segment; To manufacture a preform for a second spar segment, wherein the preform includes a sub-kink adjacent to one end of the second spar segment; Aligning the edges of the preform so that the sub-kinks are close to each other within the joint area; and Joining spar segments together within a joint region to form at least a portion of the spar detail exhibiting a kink. A method that includes this.

[0265] Clause 82. The method of joining spar segments, comprising joining spar segments in a joining region using at least one of co-curing, co-bonding, joining doubling, and fastening.

[0266] Clause 83. The method of Clause 81, wherein the manufacturing of the preform includes manufacturing the preform such that each sub-kink has an equal angular deviation.

[0267] Clause 84. The method of Clause 81, which includes manufacturing a preform such that the sub-kinks have unequal angular deviations.

[0268] Clause 85. The method of Clause 81, which includes manufacturing the preform such that the sub-kinks are fused together and the axial direction of the spur detail changes by 2 to 10 degrees.

[0269] Clause 86. Any method of Clause 81, wherein aligning the ends of the preform includes forming one of a wrap splice, a butt splice, and a scarf splice in the joint area with the ends of the preform.

[0270] Clause 87. The method of joining spar segments together, comprising applying a joining doubler to the joining region opposite the rib intersection.

[0271] Clause 88. The method of joining spar segments together, which includes inserting a sub-kink between the joining doubling and the rib.

[0272] Clause 89. The method of Clause 81, wherein joining spar segments together includes inserting a sub-kink between a front joining doubling and a rear joining doubling.

[0273] Article 90. A part of an aircraft assembled in the manner of Article 81.

[0274] Article 91. Spar details of an aircraft: A first spar segment comprising fiber-reinforced material and including a sub-kink positioned at one end; A second spar segment comprising fiber-reinforced material and having a sub-kink located at one end, wherein the end of the first spar segment having the sub-kink is adjacent to the end of the second spar segment having the sub-kink such that the sub-kinks collectively form a kink and the end defines a joining region; and A joining doubling that structurally joins the first spar segment and the second spar segment within the joining region. Super detailing with added features.

[0275] Article 92. A sub-kink associated with the first spar segment is located within approximately 30 cm (1 foot) of the end of the first spar segment; A sub-kink associated with the second spar segment is located within approximately 30 cm (1 foot) of the end of the second spar segment. Sever details of Clause 91.

[0276] Clause 93. A spar detail of Clause 92, wherein each sub-kink alters the axial direction of the spar detail by half the amount of the kink.

[0277] Clause 94. A spar detail according to Clause 91, in which sub-kinks cluster together to change the axial direction of the spar detail by 2 to 10 degrees.

[0278] Clause 95. A spar detail according to Clause 91, wherein a joining doubler is located on the opposite side of the rib intersection defined in the spar detail.

[0279] Clause 96. A spar detail of Clause 91, wherein a joining doubling is positioned to sandwich the joining area and the sub-kink.

[0280] Clause 97. For structurally joining a first spur segment and a second spur segment, a joining doubling is one of the following: co-bonded to the first spur segment and the second spur segment, co-cured, and attached by a fastener, as per the spur detail of Clause 91.

[0281] Article 98. Manufacture of an aircraft part using the spar detailing of Article 91.

[0282] Article 99. Wings of an aircraft: A first spar segment comprising fiber-reinforced material and including a sub-kink positioned at one end; A second spar segment comprising fiber-reinforced material and having a sub-kink located at one end, wherein the end of a first spar segment having the sub-kink is adjacent to the end of a second spar segment having the sub-kink such that the sub-kinks collectively form a kink and the end defines a joining region; A joining doubler that structurally joins a first spar segment and a second spar segment within the joining region; A third spar segment comprising a fiber-reinforced material, the third spar segment comprising a bonding region arranged in series with the second bonding region at the opposite end of the second spar segment; and A second joining doubling covering at least a portion of the joining region of the third spar segment and at least a portion of the second joining region of the second spar segment. The wings of an aircraft equipped with [a specific feature / feature].

[0283] Clause 100. A joining doubling is: one of the following attached to a first spar segment and a second spar segment by a fastener, the wing of an aircraft according to Clause 99: co-bonded, co-hardened, and fastened.

Claims

1. A method (700) for manufacturing a spar detail for an aircraft (10): Manufacturing (702) a preform (242) of a fiber-reinforced material (301) for spar segments (112, 114, 116), wherein at least one of the spar segments (112, 114, 116) includes a kink (370), each kink (370) is entirely contained within the preform (242) and includes a bending point (381), the bending angle between the first neutral axis (350) and the second neutral axis (350-1) of the preform (242) at the bending point (381) is between 2 and 10 degrees; To form the spar segments (112, 114, 116), solidify the preform (242) (704); and Assembling the spar segments (112, 114, 116) together (706) to form a completed spar detail (110) exhibiting at least one of the aforementioned kinks (370). A method that includes this.

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

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 the preform (242) (702) includes positioning each kink (370) at least 30 cm (1 foot) away from the end of the preform.

4. The method according to any one of claims 1 to 3, wherein each preform (242) includes at least one bonding region (341, 341-1), and assembling the spar segments (112, 114, 116) (706) includes applying bonding doublings (340, 340-1) to the spar segments (112, 114, 116) within the bonding region (341, 341-1).

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

6. Applying a joining doubler (340, 340-1) to the aforementioned spar segments (112, 114, 116) is: To form at least one of a wrap splice, a butt splice, and a scarf splice between adjacent ends of the spar segments (112, 114, 116); and Attaching the bonding doubler (340, 340-1) to the bonding region (341, 341-1) using one of co-curing, co-bonding, and fastener. The method according to claim 4 or 5, including the method described in claim 4 or 5.

7. Applying a bonding doubler (340, 340-1) to the spar segments (112, 114, 116) within the bonding region (341, 341-1) is: Position one of the joining doublers (340, 340-1) across the rib intersection (360) of the spur detail; The joining regions (341, 341-1) are sandwiched between the joining doublers (340, 340-1) and the ribs; The joining region (341, 341-1) is sandwiched between the front joining doubler and the rear joining doubler. The method according to any one of claims 4 to 6, comprising one of the following.

8. The method according to any one of claims 4 to 7, wherein the manufacturing (702) of a preform (242) of a fiber-reinforced material (301) for spar segments (112, 114, 116) comprises arranging the kink (370) so as to avoid the bonding regions (341, 341-1) of the spar segments (112, 114, 116).

9. Aircraft super detail (110): A fiber-reinforced material (301) and a first spar segment (112) including the outer end of the machine; A fiber-reinforced material (301-1), an inner end, and a second spur segment (114) including a kink with a bending point (381), wherein the bending angle between the first neutral axis (350) and the second neutral axis (350-1) of the second spur segment (114) at the bending point (381) is 2 to 10 degrees; The second spar segment (114) has an outer end of the first spar segment (112) and an inner end of the second spar segment (114) arranged in series with respect to each other to form a joint region (341); and A joining doubler (340) that covers at least a portion of the first spar segment (112) and at least a portion of the second spar segment (114) within the joining region (341), wherein the joining doubler (340) connects the first spar segment (112) and the second spar segment (114). Super detailing with added features.

10. The spar detail according to claim 9, wherein the neutral axis of the joining region (341) is aligned with the neutral axis of the first neutral axis (350) or the second neutral axis (350-1).

11. The spar detail according to claim 9 or 10, wherein the kink (370) is separated by more than 30 cm (1 foot) from the machine-side end of the second spar segment (114).

12. The spar detail according to any one of claims 9 to 11, wherein the first spar segment (112) includes a kink (370) that avoids the joining region (341).

13. The spar detail according to any one of claims 9 to 12, wherein the joining doubling (340) is located across the rib intersection (360).

14. The aforementioned joining doubler (340) is: Front joining doubler and rear joining doubler sandwiching the joining region (341); and A joining doubler (340) is positioned between the joining doubler (340) and the rib so as to sandwich the joining region (341). A spar detail according to claim 13, comprising one of the following.

15. A third spar segment (116) comprising fiber-reinforced material (301-2), an inner end, and a kink (370), wherein the inner end of the third spar segment (116) is arranged in series with the corresponding outer end of the second spar segment (114) to form a second joint region (341-1), and the kink is located outside this second joint region (341-1); and A second joining doubling (340-1) that covers at least a portion of the third spar segment (116) and at least a portion of the second spar segment (114) within the second joining region (341-1), wherein the second joining doubling (340-1) is connected to the second spar segment (114) and the third spar segment (116). A spar detail according to any one of claims 9 to 14, further comprising:

16. An aircraft wing comprising at least one spar detail as described in any one of claims 9 to 15.

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