Method, system and portion of an aircraft for manufacturing a preform

Abstract

The present application relates to a method, system and a portion of an aircraft for manufacturing a preform. A method for manufacturing a preform of a portion of an aircraft is provided, the method comprising the steps of: obtaining a sheet made of a wide-good fibrous reinforcement material; trimming the sheet to form a layup piece having a border; placing the border in alignment; setting the layup piece in a layup pattern to form a ply; performing a placement operation of transferring the layup pattern onto a layup mandrel; and shaping the layup pattern to conform to an outer shape of the layup mandrel.

Description

Technical Field

[0001] This disclosure relates to the field of manufacturing, and in particular to aircraft manufacturing. Background Technology

[0002] The mechanical structure of an aircraft is called an airframe. The airframe itself is made of discrete components such as stringers, spars, skin, and frames, which, when assembled, define the aircraft's structure. A single aircraft can be manufactured from many components. Currently, airframe components are manufactured using methods that include labor-intensive manual layup processes, or by using automated fiber placement (AFP) machines that perform layup using a single end head while traversing the shape of the airframe component to be manufactured. The airframe component remains stationary within its unit during this work.

[0003] Therefore, it is desirable to have a method and apparatus that takes into account at least some of the problems discussed above, as well as other possible problems.

[0004] The abstract of WO 2015 / 152331 states: "A prepreg sheet laminating apparatus for producing prepreg sheet laminates, wherein multiple prepreg sheets are laminated, and the prepreg sheets in one layer are laminated in a manner where the fiber orientation differs from that of the prepreg sheets in another layer. The apparatus includes: a main conveying path on which a prepreg base fabric travels; a prepreg cutting sheet forming unit that forms prepreg cut sheets by cutting the prepreg base fabric at a set angle in the width direction; and a transfer hand." The transfer hand transfers the formed prepreg cut sheet to a prepreg base fabric traveling on top of the main conveying path, and sets the prepreg cut sheet onto the prepreg base fabric, configuring the fiber direction of the prepreg cut sheet to a set direction; and a welding unit welds the prepreg base fabric and the prepreg cut sheet that has been set on the fabric.

[0005] The abstract of EP 3406431 A1 states: "Systems and methods for laying laminates are provided. One embodiment is a method comprising laying a multilayer laminate of fiber-reinforced material onto a surface by the following steps: feeding a fiber-reinforced material belt to a belt cutter that cuts the belt into segments; picking up segments of fiber-reinforced material via a pick-and-place device at each of a plurality of laminating units sequentially in the direction of travel; and placing segments of fiber-reinforced material via the pick-and-place device to form a laminate as the surface and the laminating units change position relative to each other and as the plurality of segments are laid simultaneously."

[0006] The abstract of US2009 / 148647 A1 states: "Large composite structures are manufactured by forming multiple composite laminate modules and joining these modules together along their edges using scarf joints."

[0007] Abstract of Buckingham, RO, & Newell, GC (1996). Automating the manufacturing of broadgoods. Composites Part A: Applied Science and Manufacturing, 27(3), 191-200 states: "Work entitled 'Automated handling of integrated flexible materials, design of composite components and quality assurance processes' completed under a three-year European Commission research and development contract. The paper describes the reasons for undertaking this work, the main drivers of the shaped final product, the key research areas and the technical solutions adopted, and describes, tests, and discusses the achievements and future options of the system. A construction machine is used to extract prepreg material from a standard roll, and to form multilayer hyperbolic components by removing backing sheets, visual inspection, cutting, handling, laying and consolidation. The system demonstrates that the engineering problem is not insurmountable, and that, given the right product, a semi-automated system can provide the necessary quality and productivity at an acceptable price." The abstract of EP3106280 A1 states: "A composite structure manufacturing system and method. The system includes a plurality of lamellar carriers, each lamellar carrier being configured to support at least one lamellar segment, and a forming mandrel defining an extended lamellar forming surface, the extended lamellar forming surface being shaped to define the surface profile of a composite structure. The system further includes: a carrier transfer device configured to selectively transport selected lamellar carriers from a lamellar kitting region to an intermediate position, and a forming machine configured to deform the selected lamellar carriers and corresponding lamellar segments on selected portions of the extended lamellar forming surface. The forming machine is further configured to separate the selected lamellar carriers from the corresponding lamellar segments and return the selected lamellar carriers to the carrier transfer device. These methods include methods of operating the system."

[0008] The abstract of WO 2017 / 005770 A1 states: "The present invention relates to a method for producing a fiber-metal laminate assembly of an aircraft using a manipulator system with an end effector and controls assigned to the manipulator system, wherein at least one metal layer and at least one uncured fiber layer are stacked on top of each other in a stacking sequence in a mold by the manipulator system, wherein each stacking cycle includes: picking up a metal layer or a fiber layer from a corresponding supply stack according to the stacking sequence; conveying the layer to the mold; placing the layer at a deposition surface in the mold according to the stacking sequence; and stacking the layer thus placed on the deposition surface. It is proposed that, after being picked up from the supply stack and before being stacked on the deposition surface, the layer to be stacked is deformed by the end effector so that the shape of the layer is adapted to the shape of the deposition surface."

[0009] The abstract of US2014 / 367037 A1 states: "A flexible material transfer apparatus, a flexible vacuum compaction apparatus including the flexible material transfer apparatus, a flexible vacuum chuck including the flexible vacuum compaction apparatus, and systems and methods including the like. The flexible material transfer apparatus includes a flexible substrate configured to selectively and repeatedly switch between a retracted configuration and an unfolded configuration. The flexible substrate defines a material contact surface configured to contact and be selectively and operably attached to a composite material package. The flexible substrate also defines a plurality of holding channels at least partially defined by the material contact surface and configured to have a holding vacuum applied thereto. The flexible material transfer apparatus also includes a holding manifold providing fluid communication between the plurality of holding channels and a vacuum source."

[0010] The abstract of US 2018 / 222060 A1 states: "Systems and methods for transporting material sheets are provided. One embodiment includes an apparatus for an end effector of a robot. The apparatus includes: a frame and a fixed suction cup assembly attached to the frame. The fixed suction cup assembly includes: a suction cup for holding the sheet; a pneumatic line; and a shaft coupled to the suction cup and housing the pneumatic line, the shaft enabling the suction cup to translate vertically. The apparatus also includes floating suction cup assemblies. Each floating suction cup assembly includes: a Bernoulli suction cup; a pneumatic line for applying positively pressurized gas to the Bernoulli suction cup; a shaft coupled to the Bernoulli suction cup and housing the pneumatic line, the shaft enabling the Bernoulli suction cup to translate vertically; and a bearing that enables the Bernoulli suction cup to pivot about an end of the shaft to conform to a surface." Summary of the Invention

[0011] The embodiments described herein provide an assembly line for manufacturing frame components using fiber-reinforced broad goods. The broad goods are cut, rotated, and simultaneously delivered through multiple stations to rapidly prepare lamination patterns (including one or more layers) corresponding to wing skin, fuselage skin, etc. The lamination pattern is then picked up, placed, and compacted onto a lamination mandrel. After a sufficient number of lamination patterns have been applied, the resulting preform is hardened into a composite component.

[0012] A method is provided for manufacturing a preform for a part of an aircraft. The method includes the steps of: obtaining a sheet made of a wide-width cargo fiber-reinforced material; trimming the sheet to form a laminate with boundaries; aligning the boundaries; setting the laminate according to a laminate pattern to form a layer; performing a placement operation to transfer the laminate pattern onto a laminate mandrel; and shaping the laminate pattern to conform to the shape of the laminate mandrel. Alternatively, the method includes the steps of: dividing the composite structure into zones; manufacturing laminates for each zone; setting the laminates into a laminate pattern; transferring the laminate pattern to the laminate mandrel via a carrier; and compacting the laminate pattern onto the laminate mandrel. Those skilled in the art will understand that this method can also be performed independently of the preceding methods described in this paragraph. Optionally, each laminate occupies a zone of the laminate pattern. Optionally, the method includes the step of: setting the edge of the laminate pattern onto the laminate mandrel such that the edge staggers relative to other laminate patterns on the laminate mandrel. Optionally, the step of transferring the coat pattern includes operating a pick-and-place (PNP) station. Optionally, the step of setting the coat part into a coat pattern includes setting the coat part into the shape of an wing skin. Optionally, the step of setting the coat part into a coat pattern includes setting the coat part into the shape of a fuselage skin. Optionally, the manufacturing, setting, transferring, and compacting steps are performed according to a takt time. Optionally, the takt time is synchronized with one or more feeder lines. Optionally, the takt time is different from one or more feeder lines. Optionally, the method further includes the step of compacting an additional coat pattern onto the coat pattern. Optionally, the additional coat pattern forms an interleaved splice with the coat pattern. Optionally, the method further includes the step of repeatedly performing the dividing, manufacturing, setting, transferring, and compacting steps until a preform is completed at the coat mandrel.

[0013] A forming system for manufacturing preforms for a part of an aircraft is also disclosed. The system includes: a wide-body workstation that receives sheets made of wide-body fiber-reinforced material and trims the sheets to form multiple laminates with boundaries; a rotary table that aligns the boundaries; a shuttle that holds the laminates from the rotary table in a laminated pattern; and a shuttle that transfers the laminated pattern to a laminated mandrel and shapes the pattern to conform to the shape of the laminated mandrel.

[0014] Other exemplary embodiments (e.g., methods and computer-readable media related to the foregoing embodiments) may be described below. The features, functions, and advantages already discussed may be implemented independently in different embodiments or may be combined in other embodiments, and further details of these features, functions, and advantages may be understood with reference to the following description and drawings. Attached Figure Description

[0015] Now, some embodiments of this disclosure will be described by way of example only and with reference to the accompanying drawings. In all the drawings, the same reference numerals denote the same elements or elements of the same type.

[0016] Figures 1 to 7 An assembly line for manufacturing rack assemblies from wide-width goods is illustrated in an exemplary embodiment.

[0017] Figures 7A to 7D The flow arrangement for an assembly line in an exemplary embodiment is depicted.

[0018] Figures 7E to 7F The layering pattern in the exemplary embodiments is described.

[0019] Figures 8A to 8B This is a flowchart illustrating a method for manufacturing a rack assembly from wide-width goods in an exemplary embodiment.

[0020] Figure 9 This is a diagram illustrating, in an exemplary embodiment, the division of a rack assembly into segments with uniform widths.

[0021] Figures 10 to 13 This is a diagram depicting a cutting station in an exemplary embodiment for trimming wide sheets of goods into the desired shape of a laminate.

[0022] Figure 14 A lamination cut from a wide sheet of goods is depicted in an exemplary embodiment.

[0023] Figure 15 A series of lamination patterns formed by lamination members in exemplary embodiments are depicted.

[0024] Figure 16 A pick-and-place (PNP) station is depicted in an exemplary embodiment, which places multiple coating patterns onto different circumferential portions of the coating mandrel of a half-barrel section of the fuselage skin.

[0025] Figure 17 This is a diagram illustrating the synchronization operation at the assembly line in the exemplary embodiment.

[0026] Figure 18 An assembly line for fuselage section preforms in an exemplary embodiment is depicted.

[0027] Figure 19 This is a flowchart of an aircraft manufacturing and maintenance method in an exemplary embodiment.

[0028] Figure 20 This is a block diagram of an aircraft in an exemplary embodiment. Detailed Implementation

[0029] The accompanying drawings and the following description provide specific exemplary embodiments of this disclosure. Therefore, it should be appreciated that, although not explicitly described or shown herein, those skilled in the art will be able to design various arrangements that embody the principles of this disclosure and are included within its scope. Furthermore, any examples described herein are intended to aid in understanding the principles of this disclosure and are to be construed as not being limited to such specific examples and conditions. Consequently, this disclosure is not limited to the specific embodiments or examples described below, but is limited by the claims. Composite components (such as carbon fiber reinforced polymer (CFRP) components)

[0030] Initially, it is laid out in multiple layers, collectively referred to as preforms. Individual fibers within each layer of the preform are aligned parallel to each other, but different layers exhibit different fiber orientations to increase the strength of the resulting composite component along different dimensions. The preform includes a viscous resin that solidifies to harden the preform into a composite component (e.g., for use in aircraft). Carbon fibers already impregnated with uncured thermosetting or thermoplastic resins are called "prepregs." Other types of carbon fibers include "dry fibers" that are not impregnated with thermosetting resins but may include tackifiers or adhesives. Dry fibers are foamed with resin prior to curing. For thermosetting resins, curing is a one-way process called curing, while for thermoplastic resins, if the resin is reheated, it reaches a viscous form, after which the resin can be solidified into the desired shape and solidified. As used herein, the comprehensive term for the process of transforming a preform into its final hardened shape (i.e., transforming a preform into a composite part) is “hardening,” and the term covers the curing of thermosetting preforms and the formation / solidification of thermoplastic preforms into their final desired shape.

[0031] Figures 1 to 7 An assembly line 100 for manufacturing rack assemblies from wide-width goods is illustrated in an exemplary embodiment. Figure 1 In this context, the assembly line 100 is depicted as including a wide-width goods station 110, which trims / cuts laminations 116-1 to 116-6 (collectively referred to as "laminations 116") from a roll 130 of fiber-reinforced wide-width goods material having a common width W. Specifically, the wide-width goods station 110 obtains a sheet 118 made of wide-width fiber-reinforced material 118-1 and trims the sheet 118 (e.g., applies a cut to the sheet 118) to form a plurality of laminations 116-1 to 116-6 (also collectively referred to as "laminations 116").

[0032] The sheet 118 of the wide cargo material from which the trimmed cladding 116 is derived may include a series of continuous fibers that advance along the length of the roll 130. By trimming multiple pieces from the sheet 118 and twisting these pieces to desired angles (e.g., the leading edge angle and trailing edge angle of an airfoil), cladding 116 for various fiber angles is produced.

[0033] In this embodiment, the first straight cut 113 for the plylets 116-1 to 116-6 is made from a sheet 118 of wide-width fiber-reinforced material 118-1 via a first cutting station 112. Although six plylets 116-1 to 116-6 are shown, embodiments may have more or fewer plylets (116). The number of plylets 116 can vary from sheet 175 to sheet 175. A second cutting station 114 makes a second straight cut 115 of the plylets 116 without rotating the plylets 116. The plylets 116 are then advanced to a rotary table 150 or other orientation station, where they are rotated to align 157 with a placement position 172 at the shuttle device 170. The rotary table 150 rotates the plurality of plylets 116.

[0034] The shared boundary 119 between adjacent laminations 116 is oriented at the same angle by the rotary table 150 (i.e., resulting in a butt without gaps or overlap). In one embodiment, the wide-width workstation 110 trims the sheet 118 so that the laminations 116 exhibit a shared leading edge angle 195 and a shared trailing edge angle 197, and the rotary table 150 rotates the lamination 116 by orienting the leading edge 195 of the lamination (116) to the common angle. Similarly, in another embodiment, the wide-width workstation 110 trims the sheet 118 so that the laminations 116 exhibit a shared leading edge angle 195 and a shared trailing edge angle 197, and the rotary table 150 rotates the lamination 116 by orienting the trailing edge 197 of the lamination (116) to the common angle.

[0035] The assembly line 100 sets the laminates 116 into a laminate pattern 174 for the preform 194. For example, laminates 116-1 to 116-6 can be set into laminate pattern 174, wherein the boundaries 119 of the laminates 116 are complementary to each other (i.e., aligned with each other without gaps or overlaps). Latex pattern 174 may include one or more sheets 175 formed by the laminates 116. Alternatively, completely different laminates 116 can be set into one or more sheets 175 such that when subsequent sheets 175 are added, the laminates 116 form the preform 194.

[0036] Layer patterns 174 are placed together to form a preform 194. When fully manufactured, the preform 194 includes multiple layer patterns 174 and / or individually formed layer pieces 116. Interlacing joints 177 are formed between adjacent layers 175 and / or between adjacent layer patterns 174. These interlacing joints 177 enhance the elasticity of the preform 194 after curing. Joints 177 may include butt joints, lap joints, and / or interlocking joints.

[0037] In other words, the lamination 116 and lamination pattern 174 are formed in completely different locations and transported to the preform 194 and / or lamination mandrel 190, wherein these components are butt-jointed / spliced ​​to adjacent components to facilitate the preform 194. Multiple layers 175 with staggered joints 177 are formed repeatedly in this manner until the preform 194 is fully formed, thereby producing a zonal laminate. Thus, in one embodiment, the placement operation at the assembly line 100 creates staggered joints 177 between the lamination 116 of lamination pattern 174 and the lamination 116 of another lamination pattern 174 disposed at the lamination mandrel 190.

[0038] As discussed above, in this embodiment, the laminating element 116 extends from the leading edge 195 to the trailing edge 197 of the preform 194. However, such an arrangement is not universally desired. In one embodiment, the laminating elements 116-1 to 116-6 do not coincide with the leading edge 195 or the trailing edge 197, but rather abut against adjacent laminating elements 116-1 to 116-6. In such an arrangement, the laminating elements 116-1 to 116-6 for laminating pattern 174 are not only placed side by side, but are also arranged such that they form patterns extending forward and backward, such as the leading edge 195 or the trailing edge 197 or vice versa, and a wingspan from the root at the inner end to the tip at the outer end.

[0039] The operation discussed above is for manufacturing wing skin at mandrel 190 (e.g., Figure 9 The feeder line 191 of the wing skin 900 provides input. In one embodiment, the placement of the plating element 116 and the plating pattern 174 is provided to the feeder line 191 in a Just-In-Time (JIT) manner, such that the plating pattern 174 is received within a very short period of time, such as within seconds or less than a minute while the feeder line 191 is preparing to receive the next plating pattern.

[0040] In this embodiment, in addition to the wide-piece workstation 110, the assembly line 100 also includes a small-piece workstation 120 and a tow-piece workstation 140. Depending on the implementation and design, the small-piece workstation 120 or the tow-piece workstation 140 can provide different amounts of material for the wing and can provide the majority of the material for the wing, fuselage section, or other composite structures. The small-piece workstation 120 operates a dynamic cutter workstation 122 (e.g., using a laser cutter, or a moving blade operated according to a numerical control (NC) program, a technician with hand tools, etc.) to trim and / or cut out the lamination 126 from the roll 130 of wide-piece material placed on the table 124. The size of the lamination 126 is not equal to the width of the roll 130; therefore, the lamination 126 cannot be made by two straight cuts across the width of the roll 130. The dynamic cutter workstation 122 adapts the lamination to this geometry via automated or manual lamination techniques. The tow section station 140 manufactures the cladding 148 by distributing tows 144 onto a table 146 from multiple tows 144 of fiber-reinforced material stored on rolls 142. The rolls 142 are cut to length and then kitted onto the table 146, and can then be transferred (e.g., manually or automatically) to the small parts station 120.

[0041] The coatings 116-1 to 116-6, 126, and / or 148 are conveyed from the rotary table 150 to the shuttle device 170 via a carrier 154 that slides along the frame 152. The carrier 154 operates as a placement station 155, such as an automated pick-and-place (PNP) station, an auxiliary station, or a manual station, and may, for example, include a polycarbonate thermoplastic film (e.g., LEXAN). TM A thermoplastic polycarbonate film (or other flexible material) 155-1 is used to conform the thermoplastic polycarbonate film or flexible material to the coating 116 while a vacuum is applied to controllably hold the coating 116 against the film or flexible material. The use of vacuum / suction for pick-and-place can be performed in any suitable manner known to those skilled in the art.

[0042] The shuttle 170 holds the laminating pieces from the rotary table 150 in a lamination pattern 174. Once all the laminating pieces 116 of the lamination pattern 174 have been laid, the shuttle 170 is driven along the track 160. The lamination pattern 174 forms the composite component being manufactured (such as...). Figure 9 The shape of a portion of the wing skin (900). For example, layer pattern 174 can form the shape of the wing skin (such as... Figure 9 The shape of the wing skin (900), or the shape of the fuselage skin (e.g., Figure 18(Fuselage skin 1852). In one embodiment, the cladding pattern 174 includes multiple layers made of fiber-reinforced material.

[0043] The lamination pattern 174 can be implemented as a single layer or a single sequence of laminations of a preform to be cured into a composite component, or any number of non-overlapping lamination sequences defined for the preform. In one embodiment, a technician or automated system then removes the backing from the lamination pattern 174. In other embodiments, subsequent layers are added in the manner described above before proceeding to additional operations.

[0044] Any of the operations discussed herein can be implemented in a micro-pulsed manner (where components are advanced to a length less than the length of those components and then paused), in a full-pulsed manner (where components are advanced to a length at least the length of those components and then paused), or in a continuous movement manner. In one embodiment, multiple sheets 118 of a wide-width item are used to form laminations 116-1 to 116-6 spanning from the leading edge to the trailing edge. Thus, in some embodiments, each lamination pattern 174 includes multiple rows of laminations 116-1 to 116-6.

[0045] The application pattern 174 is transferred from the shuttle 170 to the application mandrel 190 via a carrier 180 mounted to the frame 182, and the application pattern 174 is compacted into place. The carrier 180 can operate in a similar manner to the carrier 154. In one embodiment, the carrier 180 therefore operates as a pick-and-place (PNP) machine that conveys the application pattern 174 to the application mandrel 190. Figure 4 The coating pattern 174 is shaped to conform to the shape 192 of the coating mandrel 190. In one embodiment, shaping the coating pattern includes applying a vacuum (e.g., the load required to form the coating pattern) to the coating pattern 174 via a placement station 181-1 (e.g., PNP station 181) that includes a carrier 180 and a frame 182.

[0046] In another embodiment, the sheet 118 is pulsated and paused to periodically advance through the assembly line 100. Workstations on the assembly line 100 perform operations synchronously during pauses in which the laminar flow piece 116 remains stationary, and during pulsation periods in which the laminar flow piece 116 moves, synchronous operations are stopped.

[0047] In other embodiments, additional carriers 154 and 180 are used to place stacks of components such as the spars or ribs of the wing skin, the surround and window frames, and the aft pad-up onto the lamination mandrel 190 and compact those components into place.

[0048] After a sufficient number of coat patterns 174 (e.g., multiple coat patterns for each of multiple fiber orientations) have been placed and compacted onto the outline 192 of the coat mandrel 190, the preform 194 is completed. The coat mandrel 190 then leaves the assembly line 100 for further processing. A new coat mandrel 190 then replaces the previous coat mandrel 190. In one embodiment, each coat mandrel 190 includes an identifier (e.g., Figure 4 The radio frequency identifier (RFID) chip, barcode, etc. (190-1) in the coating mandrel indicates the type of composite component to which the coating mandrel 190 is intended. By reading this identifier, the controller of the assembly line 100 can confirm that the coating mandrel 190 matches the expected design.

[0049] The controller 199 typically operates the various components described above synchronously to manufacture the preform 194 of the composite part in a rapid and efficient manner. In one embodiment, the controller 199 is implemented as a custom circuit, a hardware processor that executes programmed instructions stored in memory, or a combination thereof.

[0050] In other embodiments, assembly line 100 utilizes wide-width goods of the same or similar dimensions to manufacture preforms of various components and various types of aircraft. For example, assembly line 100 may be adapted to form preforms of wing skin, spars, longitudinal spars, fuselage skin, etc. In such embodiments, the specific carrier used can be swapped out to a carrier suitable for the specific product being manufactured.

[0051] exist Figure 2 In this process, the plastering piece 116, already aligned and rotated at the rotary table 150, is conveyed via the carrier 154 to the placement position 172 at the shuttle device 170. At this point, the next plastering piece 116 is placed on the rotary table 150, and cutting 117 is performed on another plastering piece 116. Figure 3 In this process, the next lamination piece 116 has been rotated, and the carrier 154 has moved onto the turntable 150 to obtain the next lamination piece 116 and transfer it to the shuttle device 170. Although only one preform 194 of the wing skin is shown, the turntable 150 allows the lamination piece 116 to be rotated to a suitable orientation for the left and right wing skins, the upper and lower wing skins, etc. Therefore, four separate types of wing panels and / or other models of wing panels can be manufactured using a single assembly line 100.

[0052] exist Figure 4In this process, a lamination pattern 174 comprising one or more layers (such as one or two to four fiber-reinforced sheets 175) has been placed onto a shuttle device 170. Each layer is fiber-reinforced in orientation throughout the sheets and in orientation different from or the same as adjacent sheets. Each of the two to four layers must be spliced ​​to adjacent two to four layers. In one embodiment, the boundaries 119 of the lamination elements 116 in the different layers of the lamination pattern 174 are staggered, intersecting, or otherwise arranged relative to adjacent layers such that they do not cause the boundaries 119 to extend through the stack of multiple adjacent layers. The shuttle device 170 transports the lamination pattern 174 along a track 160 until the shuttle device 170 is positioned as a receiving carrier 180. Figure 5 In this process, the carrier 180 is moved above the coating pattern 174, and the coating pattern 174 is picked up via a vacuum connection. Figure 6 In this process, the carrier 180 moves along the frame 182 until the coat pattern 174 is positioned above the coat mandrel 190. The carrier 180 then releases the vacuum connection with the coat pattern 174 and applies a separate vacuum to compact the coat pattern 174 into place on the coat mandrel 190. In another embodiment, a positive airflow is provided from the carrier 180 to push the coat pattern 174 away from the carrier 180 and onto the coat mandrel 190. Figure 7 In the process, the shuttle 170 returns to the starting position 700 and, before the autoclave (not shown) or a similar device for curing, advances the lamination mandrel 190, which has received one or more lamination patterns 174 forming the preform 194, to the next process. After curing, manufacturing allowances can be partially trimmed, leaving indexing features on the remaining manufacturing allowances, or completely trimmed from the final composite part. Trimming is performed before assembly, and the precision is such that a final trimming is not required after assembly to finalize the leading and trailing edges. That is, the laminations 116-1 to 116-6 are trimmed and placed with sufficient precision so that the final laminate 175 does not require peripheral trimming to achieve the desired final panel size.

[0053] Figures 7A to 7D Flow arrangements for assembly lines in exemplary embodiments are depicted. Each of these arrangements illustrates an alternative configuration for multiple assembly lines (e.g., assembly line 100 or wide-item lamination station 151) manufacturing rack assemblies from wide-width goods. Wide-item lamination station 151 (e.g....) Figure 1 (As shown) includes: a wide-width goods station 110, a rotary table 150 or other directional stations, and placement stations 155 and 181-1 for manufacturing preforms from wide-width goods. Figures 7A to 7DAt each of the multiple wide-width cargo laminating stations, the operation is performed at the laminating mandrel.

[0054] Reference Figure 7A In the example, an assembly line 710 is depicted, comprising wide-width goods lamination stations 701-1 to 701-8 (collectively referred to as "stations 701") arranged along loop 269. Stations 701 perform operations on lamination mandrels 719, which enter via direction 718 and advance along station 701 located at loop 269 in a micro-pulsating, full-pulsating, or continuous manner. In this example, multiple mandrels 719 (such as upper wing skin panels and lower wing skin panels) are processed along loop 269 in assembly line 710. In this example, the lamination mandrel 719 progresses from left to right through horizontal segment 257 via stations 701-1, 701-2, 701-3, and 701-4, and then from right to left through stations 701-5, 701-6, 701-7, and 701-8 into the adjacent horizontal segment 263. With this left-to-right and then right-to-left arrangement through stations 701-1 to 701-8, the lamination mandrel 719 does not rotate relative to stations 701-1 to 701-8, but only shifts laterally between horizontal segment 257 and the adjacent horizontal segment 263. The preform 194 on the mandrel progresses through stations 701-1 to 701-8 until completion, and then the lamination mandrel 719 and the preform 194 exit via direction 717. Depending on the requirements, the mandrel 719 may need to pass through loop 269 and stations 701-1 to 701-8 multiple times to complete the preform 194. In some examples, one or more stations 701-n may be positioned along vertical segment 259, while in other examples, there may be no stations 701-n positioned along vertical segment 259, in which case the lamination mandrel 719 is directly displaced laterally from a horizontal segment 257 to the adjacent horizontal segment 263.

[0055] In this example, the lamination mandrel 719 reaches the end 239 of the horizontal segment 257 and shifts laterally instead of pivoting, and advances along the vertical segment 259, resulting in no change in the orientation of the lamination mandrel 719 relative to the process direction P. The lamination mandrel 719 reaches the end 239-1 of the horizontal segment 263 to complete the movement through the loop 269. The lamination mandrel 719 and the preform 194 move laterally away via direction 717 and are replaced by another lamination mandrel 719 in the loop 269.

[0056] exist Figure 7AIn this configuration, the lateral movement of the spindle 719 within the loop 269 allows for a compact assembly line 710 without additional floor space, as the spindle 719 never needs to pivot. Another advantage of this compact configuration is that the manpower required for the operating station 701 can be shared between the horizontal section 257 and the adjacent horizontal section 263.

[0057] exist Figure 7B The image depicts an assembly line 710 comprising wide-width product lamination stations 711-1 to 711-10 (collectively referred to as "stations 711") arranged along loop 712. Stations 711 perform operations on a lamination mandrel 719, which enters loop 712 via direction 718 and advances along stations 711 located on loop 712 in a micro-pulsating, full-pulsating, or continuous manner across directions 713, 714, 715, and 716. After one or more laps have passed through loop 712, the lamination mandrel 719 exits via direction 717. The orientation of the lamination mandrel 719 is rotated to align with the orientation of loop 712. Any of the stations 711 depicted herein can be implemented as described above. Figure 1 One or more of the wide-width product lamination stations 151 discussed. Stations 711-1 to 711-10 can be implemented on either side 712-1 and 712-2 of loop 712 (e.g., to facilitate return loop operation), and lamination mandrel 719 can receive laminations from one or more of the stations 711. Laminations (e.g., lamination piece 116) can be spliced ​​together to form a sheet 175, and the splicing positions can be staggered from sheet to sheet to avoid stacking splices directly on top of previous splices. Lamination mandrel 719 continues to receive lamination piece 116 until preform 194 is completed. In one embodiment, lamination piece 116 exhibits a different orientation. After preform 194 has received all sheets, preform 194 and mandrel 719 leave for further processing, and one or more other lamination mandrels 719 are added to loop 712. In one embodiment, multiple spindles 719 advance synchronously along a loop 712, wherein these components pulsate and pause simultaneously. The orientation of the loop 722 may be the process direction P.

[0058] exist Figure 7CThe image depicts an assembly line 720 comprising wide-width product lamination stations 721-1 to 721-3, and 721-6 to 721-10 (collectively referred to as "stations 721") arranged along a loop 722 forming a "U" shape 722-3. Stations 721 perform operations on a lamination mandrel 729, which enters the loop 722 via direction 728 and advances along stations 721 located at the loop 722 in a micro-pulsating, full-pulsating, or continuous manner across directions 725, 726, and 723. The direction of the lamination mandrel 729 is rotated to align with the orientation of the loop 722. After advancing along the loop 722, the lamination mandrel 729 and the preform 194 exit via direction 727. Any of the stations 721 depicted herein can be implemented as shown above. Figure 1 One or more of the wide-width product lamination stations 151 discussed. Stations 721-1 to 721-3 and stations 721-6 to 721-10 can be implemented on each side 722-1 and 722-2 of loop 722, and lamination mandrel 729 can receive laminations from one or more of the stations 721. Laminations (e.g., lamination piece 116) can be spliced ​​together to form a sheet 175, and the splicing positions can be staggered from sheet to sheet to avoid stacking splices directly on top of previous splices. Lamination mandrel 729 continues to receive lamination piece 116 until preform 194 is completed. In one embodiment, lamination piece 116 exhibits a different orientation. After preform 194 has received all the sheets, preform 194 and mandrel 729 leave for further processing, and one or more other lamination mandrels 729 are added to loop 722. In one implementation, multiple spindles 729 advance synchronously along a ring 722, wherein these components pulsate and pause simultaneously.

[0059] exist Figure 7D The image depicts an assembly line 730 comprising wide-width product lamination stations 731-1 to 731-5 (collectively referred to as "stations 731") arranged along a loop 732 forming an "S" shape 732-3. Stations 731 perform operations on a lamination mandrel 739, which enters the loop 732 via direction 738 and advances along stations 731 located at the loop 732 in a micro-pulsating, full-pulsating, or continuous manner, across directions 733, 734, 735, and 736. After advancing through the loop 732, the lamination mandrel 739 exits via direction 737. Any of the stations 731 depicted herein can be implemented as described above. Figure 1One or more of the wide-width product lamination stations 151 discussed herein. Stations 731-1 to 731-10 may be implemented on each side 732-1 and 732-2 of loop 732, and lamination mandrel 739 may receive laminations from one or more of the stations 731. Laminations (e.g., lamination piece 116) may be spliced ​​together to form a sheet 175, and the splicing positions may be staggered from sheet to sheet to avoid stacking splices directly on top of previous splices. Lamination mandrel 739 continues to receive lamination piece 116 until preform 194 is completed. In one embodiment, lamination piece 116 exhibits a different orientation. After preform 194 has received all sheets, preform 194 and mandrel 739 depart for further processing, and one or more other lamination mandrels 739 are added to loop 732. In one implementation, multiple spindles 739 advance synchronously along a ring 732, wherein these components pulsate and pause simultaneously.

[0060] Figures 7E to 7F The layering pattern in the exemplary embodiments is described. Figure 7E This is a top view of the lamination pattern 740 of the laminate 749 in the exemplary embodiment. According to... Figure 7E The lamination pattern 740 includes zones A and B, as indicated by "A" and "B" on the lamination 116 of the laminate 749. The splices 741 between the laminations 116 in zones A and B vary along the length L of the laminate 749, thereby forming an interlaced pattern 743 and preventing the formation of a single seam along the length of the lamination map. That is, zone lamination is performed such that the boundaries 745 between the laminations 116 in zones A and B are staggered across layers to avoid overlapping of the boundaries 745 and / or splices 741. In other embodiments, the zones are overlapped in an angled shape according to the fiber orientation of the material being laid.

[0061] Each splice 741 can be formed by placing different laminations 116. Although the splices 741 are shown as lines, each splice 741 occupies a narrow area between adjacent zones, where the laminations 116 from these zones are butt-jointed, overlapped, or otherwise physically integral with each other. Each layer 175 being spliced ​​can have a unique boundary 745, and the boundary 745 can vary by a fraction of a centimeter (or inch) between adjacent layers, resulting in a thickness-staggered splice 741 that runs through the zones. That is, the position of the splices 741 changes incrementally between layers, thereby forming an interlaced pattern 743 (e.g., a stepped pattern, an interlaced shape, etc.) that runs through multiple layers. The interlacing prevents overlapping on top of each other and causing material build-up. Therefore, in one embodiment, the cutting position for the splice 741 changes between the layers 175. The splice 741 extends across multiple layers. In this embodiment, the splices 741 are selected / placed such that they do not intersect with the supports 742, in order to prevent a significant increase in thickness or complexity near the supports 742. Therefore, the boundaries 745 are staggered from sheet 175 to sheet 175.

[0062] Figure 7F A similar arrangement of the laminating element 116 in zones A and B is illustrated, as well as the splicing 741 of the wing skin lamination 750. The lamination 750 also includes a support 742, which can be used to provide reinforcement for access panels, flaps, etc. Figures 7E to 7F This illustrates how the zone lamination technique discussed in this paper can be used in a variety of laminate designs.

[0063] Regarding Figure 8A Let us now discuss exemplary details of the operation of assembly line 100. For this embodiment, it is assumed that a roll 130 of a wide product comprising continuous fiber-reinforced material has been loaded onto assembly line 100.

[0064] It should be noted that the term "mandrel" as used in this application refers to a mandrel on which components (e.g., aircraft components) can be positioned. The term mandrel is used interchangeably with the term "coating mandrel." A mandrel is provided with and / or used as a surface on which components, material layers, or combinations thereof can be positioned. This can be, for example, a coating mandrel or a mandrel on which wing panels and / or wing skin are transferred. It will be apparent to those skilled in the art that the coating mandrels 190, 719 illustrated in the accompanying drawings of this disclosure are schematic illustrations of coating mandrels 190, 719, which may have a certain 3D shape, structure, and / or surface as described above.

[0065] Figure 8AThis is a flowchart illustrating a method 800 for manufacturing a frame in an exemplary embodiment. (See also...) Figure 1 The steps of method 800 are described using assembly line 100 as an example; however, those skilled in the art will recognize that method 800 can be performed 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 also be performed in an alternative order.

[0066] In step 802, a sheet 118 of fiber-reinforced wide goods is obtained from the assembly line 100. In one embodiment, the step of obtaining the sheet 118 includes threading the sheet 118 through a first cutting table 112 and a second cutting table 114.

[0067] Step 804 includes trimming sheet 118 (e.g., applying cuts to sheet 118) to form a laminate 116 having complementary boundaries 119. The cuts to each laminate in the laminate 116 are applied as a first straight cut 113 from a first cutting station 112, followed by a second straight cut 115 from a second cutting station 114. Each first straight cut 113 is made across the entire width of the sheet 118 of the wide cargo at a desired angle (e.g., a leading edge angle 195° or a trailing edge angle 197°). At this point, the second straight cut 115 can be applied to the segment produced by the first cuts. For example, for each segment, one cut may correspond to a leading edge angle 195° on a wing skin, while another cut may correspond to a trailing edge angle 197° on a wing skin. This allows for the cutting of a constant-width wide cargo material via two cuts to produce segments with two parallel edges, both cut into a tapered shape to form a desired shape for placement as, for example, a layer of wing skin. In short, the composite structure is formed by several laminates 116, which are produced by two straight cuts from a wide cargo piece and then placed adjacent to each other. The arrangement of the cuts and the sizing of the laminates 116 result in little or no material waste, thereby reducing the "buy to fly" cost of the final aircraft. This method of using straight cuts saves time and eliminates complexity while also reducing waste.

[0068] Step 806 includes aligning the boundaries 119 (e.g., by rotating and translating the laminar flow 116). In this embodiment, because the edges of the sheet 118 of the wide-width goods form the sides of the laminar flow 116, the laminar flow 116 can be placed side-by-side without overlap by placing the boundaries 119 together (i.e., by butt-joining the laminar flow 116 as part of a splicing process). In this embodiment, the edges of the sheet 118 of the wide-width goods form complementary sides of the laminar flow 116. By aligning these complementary sides in a manner that rotates them to a common angle, the laminar flow 116 can be easily combined into a single laminar flow pattern 174 by translating it. In one embodiment, trimming the sheet / applying a cut to the sheet 118 produces laminar flow 116 exhibiting a common leading edge angle 195 and a common trailing edge angle 197, and the step of rotating the laminar flow 116 includes orienting the leading edge of the laminar flow 116 to a common angle (e.g., a leading edge angle or a trailing edge angle). In other embodiments, a single coat pattern 174 is formed by using several rows of coat pieces 116-1 to 116-6 spanning the chordal direction.

[0069] Step 808 includes: setting the laminating piece 116 onto the shuttle device 170 according to the laminating pattern 174 to form one or more sheets 175. That is, the laminating pattern 174 itself includes one or more sheets 175. In this embodiment, the step includes: laterally moving the shuttle device 170 until a new placement position 172 is exposed, and then moving the carrier 154 until the laminating piece 116 is aligned with the placement position 172. Then, the laminating piece 116 is placed in place. These operations continue for multiple laminating pieces 116 until the entire laminating pattern 174 is produced. According to an embodiment, the step of setting the laminating piece 116 into the laminating pattern 174 includes: setting the laminating piece 116 into a wing skin (e.g., Figure 9 The shape of the wing skin (900) or the fuselage skin (e.g., Figure 18 The shape of the fuselage skin 1852). Thus, in one embodiment, the layer pattern 174 forms the wing skin (e.g., Figure 9 The wing skin 900) layer 175. In another embodiment, the step of setting the cladding 116 as cladding pattern 174 includes: setting the cladding 116 as fuselage skin (e.g., Figure 18 The shape of one or more layers 175 of the fuselage skin (1852).

[0070] Step 810 includes performing a placement operation to transfer the application pattern 174 onto the application mandrel 190. This operation includes moving the shuttle 170 along the track 160, then picking up the entire application pattern 174 at once using the carrier 180, and moving the application pattern 174 to the application mandrel 190. In other embodiments, placement is performed manually rather than via an automated pick-and-place (PNP) process. Thus, in one embodiment, the PNP process is performed automatically via a PNP station 181, while in another embodiment, placement is performed manually.

[0071] In step 812, the coat pattern 174 is shaped to conform to the shape 192 of the coat mandrel 190. This includes driving the carrier 180 into the coat mandrel 190 to conform the coat pattern 174 to the shape 192 of the coat mandrel 190. In one embodiment, shaping the coat pattern 174 includes applying a vacuum to the coat pattern 174 via the carrier 180 to consolidate / fit / compact the coat pattern during the PNP operation. Thus, in one embodiment, the operation of shaping the coat pattern is performed by the carrier 180 for performing the placement operation.

[0072] The steps of method 800, including trimming, placement, setting, performing placement operations, and forming, can be repeated until the preform 194 is completed at the coating mandrel 190. Moreover, in one embodiment, the operations of acquiring, trimming / applying cutting, rotating, setting, placing (e.g., performing PNP operations), and forming are performed synchronously in a pulsed manner via stations of the assembly line 100, wherein each pulse is followed by a pause.

[0073] In other embodiments, the operations of acquiring, trimming / applying cutting, rotating, setting, placing (e.g., performing PNP operations), and forming are performed according to cycle time. In such embodiments, assembly line 100 operates as a feeder line, wherein the cycle time of assembly line 100 is synchronized with or different from the cycle time of one or more other feeder lines (e.g., feeder line 191 for manufacturing wing skin at mandrel 190). Multiple feeder lines such as assembly line 100 can provide multiple plating pieces 116 along an annular, "S"-shaped, or "C"-shaped assembly line (e.g., ...). Figures 7A to 7D The system can be used at multiple locations to apply multiple sublayer sheets 175 to the coating mandrel and / or preform as the coating mandrel 190 and / or preform progresses along the processing direction. The cycle time of the feeder line does not need to be compared with that used in assembly lines that are circular, "S" shaped, or "C" shaped (e.g., Figures 7A to 7DThe cycle time of the systems is the same. In one embodiment, multiple assembly lines 100 simultaneously supply the cladding 116 to the same wing skin.

[0074] Moreover, although assembly line 100 is described as manufacturing wing skin, in other embodiments, assembly line 100 is used to manufacture fuselage sections, tail sections, engine nacelles, doors, flaps, slats and / or other components.

[0075] Figure 8B This is a flowchart illustrating a method 850 for manufacturing a rack composite structure in an exemplary embodiment. Step 852 includes: dividing the composite structure into multiple zones (e.g., Figure 9 Zones 902-2 to 902-5, collectively referred to as zone 902). In one embodiment, the zone (e.g., Figure 9 The zones 902-2 to 902-5 correspond to the shape and size of placement position 172.

[0076] Step 854 includes manufacturing the coat pieces 116 for each zone in zone 902. Manufacturing the coat pieces 116 includes performing the trimming / cutting operations discussed above at a station of assembly line 100. In step 856, the coat pieces 116 are configured as coat pattern 174. Each coat piece 116 occupies a zone of coat pattern 174 (e.g., Figure 9 (Any of zones 902-2 to 902-5). This includes: operating the rotary table 150 and the carrier 154 to place the cladding 116 into the cladding pattern 174. In one embodiment, the step of setting the cladding 116 into the cladding pattern 174 includes: setting the cladding 116 into a wing skin (e.g., Figure 9 The shape of the wing skin (900) or the fuselage skin (e.g., Figure 18 The shape of the fuselage skin (1852).

[0077] Step 858 includes: transferring the application pattern to the application mandrel 190 via a carrier. In this embodiment, this operation is performed by advancing the shuttle 170 carrying the application pattern 174 below the carrier 180. Thus, in one embodiment, the step of transferring the application pattern 174 includes operating the pick-and-place (PNP) station 181.

[0078] Step 860 includes: compacting the coat pattern 174 onto the coat mandrel 190 to manufacture a preform 194. In this embodiment, this includes picking up, placing, and compacting the entire coat pattern 174 onto the coat mandrel 190 in one go using a carrier 180. In another embodiment, the method further includes the step of: positioning an edge of the coat pattern 174 onto the coat mandrel 190 such that the edge is staggered relative to other coat patterns 174 on the coat mandrel 190.

[0079] Methods 800 and 850 offer significant advantages over the prior art because they enable the rapid manufacture of large composite structures from wide-width goods without slow trimming / cutting processes, such as manual or automated processes that use a single cutter tip to trace the shape of the composite component. In other embodiments, there may be more than one wide-width goods station 110 or small-parts station 120 that simultaneously feeds material to an assembly line arranged in a circular, "S," or "C" shape (e.g., ...). Figures 7A to 7D The system has multiple placement stations. This type of parallel processing accelerates the manufacturing process. Moreover, since pick-and-place operations are much faster than using placement ends to lay a piece and then trim it, using wide-width goods also increases the speed of laying.

[0080] Figure 9 This diagram illustrates, in an exemplary embodiment, the division of a frame assembly into segments of uniform width. In this embodiment, the frame assembly includes a wing skin 900. The wing skin 900 is subdivided into strips 902-2 to 902-5 cut from a wide piece of material having a uniform width W. The final strip 904 for a small lamination and the first strip 901 for another lamination have been trimmed / cut multiple times from the wide piece of material to achieve the desired dimensions. In this case, W corresponds to the width of the wide piece of material sheet.

[0081] Figures 10 to 13 This is a diagram depicting a wide-width goods handling station (1000). The wide-width goods handling station (1000) is... Figures 1 to 7Example wide-format goods station 110. In an exemplary embodiment, wide-format goods station 1000 includes a cutting station for trimming / cutting sheets 118 of wide-format goods into desired shapes for laminates. In this embodiment, a first cutting station 1012 includes a rotating component 1022 on which a cutter 1032 is mounted. The rotating component 1022 may have a smooth surface that prevents twisting, wrinkling, and / or bunching of the sheet 118. The cutter 1032 is aligned to cut at a single desired angle and through the sheet 118 of the wide-format goods material in a straight cut. Therefore, the cutter 1032 is long enough to accommodate the desired cutting angle. A second cutting station 1010 operates a cutter 1040 to perform a second cut. The second cutting station 1010 includes a rail 1030 on which the cutter 1040 moves back and forth.

[0082] exist Figure 11 In this process, the sheet 1100 of the wide-width goods material 1102 is advanced along the direction 1090 through the first cutting station 1012, and the cutter 1032 performs the first cut 1110. Figure 12 In this process, the additional sheet 1100 is fed to the first cutting station 1012, where an additional first cut 1110 is made (this first cut may or may not be at the same angle as the previous first cut, depending on the design). During the movement of the sheet 1100 along direction 1090, the second cutting station 1010 slides the cutter 1040 along guide rail 1030 along direction 1290, resulting in an angled cut 1210 at the plating 1200. In other embodiments, additional cutting operations are performed to produce a polygonal plating with any number of desired sides and angles. This produces... Figure 13 The new coating element 1300 is depicted.

[0083] Figure 14The illustration depicts laminations 1420-1 to 1420-6 (collectively or individually referred to as "laminations 1420") cut from a sheet 1400 of a wide-width garment, disposed on surface 1430 in an exemplary embodiment. In this embodiment, each lamination 1420 is cut to exhibit a leading edge angle θ1 and a trailing edge angle θ2, and each lamination 1420 is cut from the same sheet (or sheet of the same width) of the fiber-reinforced wide-width garment. Some of the laminations 1420 receive additional cuts 1490 to obtain desired dimensions. Furthermore, although not shown, the laminations 1420 are cut with a flat pattern, wherein the laminations 1420 are butted and abutted against each other, and the flat pattern is stacked onto a mandrel (e.g., having a complex profile) to obtain a complex shape. That is, the laminate 1420 is trimmed in such a way that it conforms to the complex shape of the flat pattern while still mating adjacent laminates 1420 for easy splicing. After the laminates 1420 are cut via the cutting station 1410, they are seamlessly conveyed and aligned / rotated into a laminate pattern 1460 with uniform trailing edges 1440 and leading edges 1450 (as shown by arrows 1470-1 to 1470-6), and mated to adjacent pieces. This mating operation is part of the splicing when the mating is staggered from lamination 175 to lamination 175 by stacking (i.e., layers through the preform 194).

[0084] Waste material 1492 can be recycled or discarded as needed, or used in another location as another layer 1420 at another location. In this depiction, for ease of illustration, the layer pattern 1460 is shown in an exploded view. The layer 1420 is positioned as a horizontal and vertical mirror image of adjacent pieces, or at 180 degrees relative to each other, so that a common cutting angle can be achieved between the pieces. That is, the cutting of a piece according to angle θ2 is also according to angle θ.

[0085] 2. Cut adjacent parts. The sum of the leading edge angle θ1 and the trailing edge angle θ2 may be equal to 180° or not equal to 180°. When the sum of the leading edge angle θ1 and the trailing edge angle θ2 is not equal to 180°, adjacent parts are arranged at 180 degrees relative to each other. In this way, the laminating part 1420 can be used efficiently to form a wing skin with minimal waste 1492. In another embodiment, one laminating part 1420 is assembled into a right or upper panel layout, while another laminating part 1420 is assembled into a left or lower panel layout. These arrangements can provide particular benefits in systems where stations are dedicated to the specific size and / or orientation of the laminating part 1420, and provide roughly similar laminating parts for each wing skin. Systems like this can also be implemented for setups with multiple laying positions around an assembly line in a ring, "S" shape, or "C" shape (e.g., Figures 7A to 7D (system).

[0086] Figure 15 A series of lamination patterns formed from laminations in an exemplary embodiment are depicted. For the same wing skin preform 1500, lamination pattern 1510 exhibits a +45° fiber orientation 1512, lamination pattern 1520 exhibits a 90° fiber orientation 1522, lamination pattern 1530 exhibits a -45° fiber orientation 1532, and lamination pattern 1540 exhibits a 0° fiber orientation 1542. In each of 1510, 1520, 1530, and 1540, the fiber orientation is referenced to the orientation of a wide sheet of material having a width W (e.g., 118, 1100, or 1400). By utilizing different fiber orientations for different layers of the wing skin, a desired level of structural strength is achieved.

[0087] Figure 16 This is a top view depicting a PNP (Plug-in-Place) station in an exemplary embodiment, which places multiple plating patterns onto different circumferential portions of the plating mandrel 1610 of the half-barrel segment of the fuselage skin. Although the foregoing figures depict wings, flaps, stabilizers, etc., similar systems can be used for the fuselage, engine nacelle, doors, or other structural components of an aircraft. Therefore, Figure 16This paper describes one of many possible wide-width cargo PNP techniques for strip lamination of arched sections of fuselage skin. In this embodiment, a carrier 1620 moves along a frame 1622. The carrier 1620 utilizes a vacuum connection to hold the lamination pattern 1630 of the laminated parts. The carrier 1620 picks up and places individual strips during a single PNP operation, and then compacts the strip. For example, the carrier 1620 first picks up the lamination pattern 1630 for the left strip 1612 of the lamination mandrel 1610, then picks up the lamination pattern 1630 for the top strip 1614 of the lamination mandrel 1610, and then picks up the lamination pattern 1630 for the right strip 1616 of the lamination mandrel 1610. By repeatedly performing these operations and staggering the positions of the lamination patterns 1630 relative to the lamination patterns in other layers of the fuselage skin (i.e., the positions where the patterns are aligned with each other), the resulting fuselage skin exhibits the desired structural strength. Specifically, for lamination patterns exhibiting + / -45° or 90° fiber orientation, and even for lamination patterns with the same orientation placed above each other, the lamination patterns are designed such that the boundaries of the laminations in different layers overlap or form butt joints. In other embodiments, the PNP station 1600 uses the techniques discussed above to place the door frame or similar support at the location of the fuselage or wing skin. For example, Figure 16 The diagram also depicts a PNP station 1640 for the door frame. The PNP station 1640 includes a carrier 1642 that moves along a frame 1622. A plating pattern 1650 (or plating package, or other component) for the door frame is held at the carrier for placement on a zone of the plating mandrel 1610.

[0088] Figure 17 This is a flowchart 1700 illustrating the synchronization operation at the assembly line in the exemplary embodiment. Figure 17Examples illustrate how trimming / cutting can be synchronized across workstations, coordinating trimming / cutting operations, rotation at the turntable, and movement of the shuttle with cycle times (e.g., desired production times for wing or fuselage skin), and allowing workstations to perform operations synchronously. For instance, a first cut 1701 can be applied at a first cutting table 112 while a second cut 1702 is being applied at a second cutting table 114. Further operations, such as rotation 1703 and transfer via shuttle 1704, can be performed while the first cut 1701 and the second cut 1702 are being executed. Similarly, workstation 1722 of lamination workstation 1720 can perform lamination 1707, and workstation 1724 can simultaneously perform lamination 1707 on different lamination pieces. Further operations, such as applying the lamination to mandrel 1705, returning to mandrel 1706, and / or transferring the lamination piece to shuttle 1709, can also be performed synchronously. That is, trimming / cutting operations, shuttle movement, and rotary table rotation can occur during pauses between pulsations of wide-width sheet goods moving from rolls. In one implementation, the pulsation and accompanying pauses last for several seconds. This insight increases work density and throughput while ensuring that each station does not interfere with the operation of other stations.

[0089] Figure 18An assembly line 1800 for preformed fuselage sections is depicted in an exemplary embodiment. Assembly line 1800 includes wide-section production lines 1810-1 to 1810-3 that trim / cut laminations 1830-1 to 1830-3 for placement in lamination pattern 1851 of fuselage skin 1852. Wide-section production lines 1810-1 to 1810-3 include rotary tables 1820-1 to 1820-3. Rotary tables 1820-1 to 1820-3 rotate laminations 1830-1 to 1830-3 to a desired alignment, and a carrier 1821 applies laminations 1830-1 to 1830-3 to conveyor devices 1840-1 to 1840-3. The laminating element forms a laminating pattern, which is transferred via carriers 1850-1 to 1850-3 to zones 1860-1 to 1860-3 of the laminating mandrel 1870 as laminating pattern 1851. The door frame station 1880 manufactures a preform 1882 of the door frame for application to the laminating mandrel 1870 via carrier 1850-1. The preform 1882 is manufactured from a fiber-reinforced preform 1816, which is laid at the lamination station 1814. The system 1800 can also be adapted to be placed on the laminating mandrel 1870, which is longer than the carriers 1850-1 to 1850-3. In this case, the coating mandrel 870 can be moved laterally to allow the coating pieces (116) 1830-1 to 1830-3 to be placed multiple times in the various zones 1860-1 to 1860-3.

[0090] In the following examples, additional processes, systems, and methods are described in the context of an assembly line for manufacturing composite components from wide-width goods.

[0091] Referring more specifically to the accompanying drawings, embodiments of this disclosure can be implemented as follows: Figure 19 The aircraft manufacturing and maintenance methods shown in 1900 and as follows Figure 20The description is presented within the context of the aircraft 1902. During pre-production, method 1900 may include the specification and design 1904 of aircraft 1902 and material procurement 1906. During production, the manufacturing of components and sub-components of aircraft 1902 and system integration 1910 may be carried out. Subsequently, aircraft 1902 may undergo certification and delivery 1912 for use 1914. When in use by the customer, routine maintenance and upkeep 1916 are performed on aircraft 1902 (this may also include modifications, reconfigurations, refurbishments, etc.). The equipment and methods specifically implemented herein may be employed during any one or more suitable stages of production and use as described in Method 1900 (e.g., Specification and Design 1904, Material Procurement 1906, Component and Sub-component Manufacturing 1908, System Integration 1910, Certification and Delivery 1912, In Use 1914, Maintenance and Care 1916) and / or any suitable component of Aircraft 1902 (e.g., Frame 1918, System 1920, Interior 1922, Propulsion System 1924, Electrical System 1926, Hydraulic System 1928, Environment 1930).

[0092] Each process in Method 1900 can be performed or implemented by a system integrator, a third party, and / or an operator (e.g., a customer). For the purposes of this description, a system integrator can include, but is not limited to, any number of aircraft manufacturers and main system subcontractors; a third party can include, but is not limited to, any number of vendors, subcontractors, and suppliers; and an operator can be an airline, leasing company, military entity, service organization, etc.

[0093] like Figure 20 As shown, an aircraft 1902 produced according to method 1900 may include a frame 1918 having multiple systems 1920 and an interior 1922. Examples of systems 1920 include one or more of the following: a propulsion system 1924, an electrical system 1926, a hydraulic system 1928, and an environmental system 1930. Any number of other systems may be included. Although an aerospace example is shown, the principles of the invention can be applied to other industries such as the automotive industry.

[0094] As mentioned above, the equipment and methods specifically implemented herein can be used during any or more suitable stages of the production and maintenance phases described in method 1900. For example, the components or sub-components corresponding to component and sub-component manufacturing 1908 can be made or manufactured in a manner similar to that of components or sub-components produced when aircraft 1902 is in use. Moreover, during sub-component manufacturing 1908 and system integration 1910, one or more equipment implementations, method implementations, or combinations thereof can be utilized, for example, by significantly accelerating the assembly of aircraft 1902 or reducing the cost of the aircraft. Similarly, when aircraft 1902 is in use (e.g., and without limitation, during maintenance and servicing 1916), one or more equipment implementations, method implementations, or combinations thereof can be utilized. Therefore, the present invention can be used at any stage or in any combination thereof discussed herein, such as specifications and design 1904, material procurement 1906, component and sub-component manufacturing 1908, system integration 1910, certification and delivery 1912, in use 1914, maintenance and servicing 1916, and / or any suitable component of the aircraft 1902 (e.g., frame 1918, system 1920, interior 1922, propulsion system 1924, electrical system 1926, hydraulic system 1928 and / or environment 1930).

[0095] In one embodiment, the component comprises a portion of a frame 1918 and is manufactured during component and sub-component manufacturing 1908. The component can then be assembled into the aircraft during system integration 1910 and utilized during use 1914 until wear renders it unusable. Then, during maintenance 1916, the component can be discarded and replaced with a newly manufactured component. The inventive components and methods can be readily utilized throughout component and sub-component manufacturing 1908 to produce new components.

[0096] Any of the various control elements (e.g., electrical or electronic components) shown in the figures or described herein can be implemented as hardware, a processor executing software, a processor executing firmware, or a combination thereof. For example, an element can be implemented as dedicated hardware. The dedicated hardware element can be referred to as a “processor,” a “controller,” or a similar term. When provided by a processor, the functionality can be provided by a single dedicated processor, a single shared processor, or multiple individual processors, some of which may share the functionality. Furthermore, the terms “processor” or “controller” as explicitly used should not be construed as referring specifically to hardware capable of executing software, but may implicitly include, but are not limited to, digital signal processor (DSP) hardware, network processors, application-specific integrated circuits (ASICs) or other circuitry, field-programmable gate arrays (FPGAs), read-only memory (ROM) storing software, random access memory (RAM), non-volatile memory, logic, or some other physical hardware component or module.

[0097] Furthermore, control elements can be implemented as instructions executable by a processor or computer to perform the function of that element. Some examples of instructions are software, program code, and firmware. The instructions are operable when executed by a processor to instruct the processor to perform the function of the element. The instructions can be stored on a storage device that can be read by a processor. Some examples of storage devices are digital or solid-state memories, magnetic storage media such as disks and tapes, hard disk drives, or optically readable digital data storage media.

[0098] The following paragraphs describe illustrative, non-exclusive examples of background techniques that are useful for understanding the present invention.

[0099] According to an aspect of this disclosure, a method (800) for manufacturing a preform (194) for a part of an aircraft (1902) is disclosed, the method comprising the following steps:

[0100] Obtain (802) a sheet (118) made of wide-width cargo fiber-reinforced material;

[0101] Trim (804) the sheet (118) to form a cladding (116) with a boundary (119);

[0102] Position the boundary (119) (806) for alignment;

[0103] The laminating element (116) is set according to the laminating pattern (174) to form a sheet (175);

[0104] Perform (810) a placement operation to transfer the plating pattern (174) onto the plating mandrel (190); and

[0105] The coating pattern (174) is shaped (812) to match the shape (192) of the coating mandrel (190).

[0106] Optionally, the boundaries (119) of the cladding (116) are complementary to each other.

[0107] Optionally, the placement operation described in (810) forms an interlocking joint between the plastering piece (116) of the plastering pattern (174) and the plastering piece (116) of another plastering pattern (174) disposed at the plastering mandrel (190).

[0108] Optionally, trimming (804) the sheet (118) produces the lamination (116) exhibiting a common leading edge angle and a common trailing edge angle, and the laminations (116) are mated together when arranged according to the lamination pattern (174). Optionally, the step of aligning the boundary (119) includes rotating the lamination (116) by orienting the leading edges of the laminations (116) to a common angle.

[0109] Optionally, the step of positioning (806) the boundary (119) for alignment includes rotating the drape (116) by orienting the trailing edge of the drape (116) to a common angle.

[0110] Optionally, each coating element (116) occupies a zone (902) of the coating pattern (119).

[0111] Optionally, the layering pattern (174) forms a layer (175) of the wing skin (900).

[0112] Optionally, the step of forming (812) the coating pattern (174) includes applying a vacuum to the coating pattern via a carrier (180) when the placement (806) operation is performed.

[0113] Optionally, the step of setting the cladding member (116) (808) into the cladding pattern (174) includes setting the cladding member (116) into the shape of at least one layer (175) of the wing skin (900).

[0114] Optionally, the step of setting the laminating member (116) (808) into the laminating pattern (174) includes setting the laminating member (116) into the shape of at least one layer (175) of the fuselage skin (1852).

[0115] Optionally, the operations of acquiring (802), trimming (804), placing (806), setting (808), executing (810), and forming (812) are performed via a station of an assembly line (100) that operates synchronously in a pulsed manner, wherein a pause follows an operation pulse.

[0116] Optionally, the operations of acquiring (802), trimming (804), placing (806), setting (808), executing (810), and shaping (812) are performed according to a cycle time. Optionally, the cycle time is synchronized with one or more feeder lines (191).

[0117] Optionally, the cycle time is different from that of one or more feeder lines (191).

[0118] Optionally, the step of placing the layer pattern (174) described in (810) is performed in a just-in-time (JIT) manner.

[0119] Optionally, before performing the placement operation described in (810), an additional layer pattern (174) is set (808) on the layer pattern (174).

[0120] Optionally, the method further includes the following steps:

[0121] The process of trimming (804), placing (806), setting (808), executing (810), and shaping (812) is repeated until the preform (194) is completed at the coating mandrel (190).

[0122] Optionally, the step of forming (812) the coating pattern (174) is performed by a carrier (180) for performing the placement operation (810).

[0123] Optionally, the method further includes the following steps:

[0124] The composite structure is divided into zones (902) (852); and

[0125] Fabricate (854) the coating elements (116) of each zone in the zone (902). Optionally, each coating element (116) occupies a zone (902) of the coating pattern (174).

[0126] Optionally, the method further includes the following steps:

[0127] The edges of the coating pattern (174) are set on the coating mandrel (190) such that the edges are staggered relative to other coating patterns (174) of the coating mandrel (190).

[0128] Optionally, the step of transmitting (858) the coating pattern (174) includes: operating the pick-and-place (PNP) station (181).

[0129] Optionally, the step of setting (856) the coating member (116) to the coating pattern (174) includes:

[0130] The cladding component (116) is configured in the shape of a wing skin (900).

[0131] Optionally, the step of setting (856) the coating member (116) to the coating pattern (174) includes:

[0132] The cladding (116) is configured in the shape of a fuselage skin (1852).

[0133] Optionally, the steps of manufacturing (854), setting (856), conveying (858), and compacting (860) are performed according to a cycle time. Optionally, the cycle time is synchronized with one or more feeder lines (191). Optionally, the cycle time is different from that of one or more feeder lines (191).

[0134] Optionally, the method further includes the following steps:

[0135] The additional coating pattern (174) is compacted (860) onto the coating pattern (174). Optionally, the additional coating pattern (174) and the coating pattern (174) are staggered.

[0136] Optionally, the method further includes the following steps:

[0137] The steps of dividing (852), manufacturing (854), setting (856), conveying (858) and compacting (860) are repeatedly performed until the preform (194) is completed at the coating mandrel (190).

[0138] According to aspects of this disclosure, a portion of an aircraft (1902) assembled according to the method (800) of any of the foregoing examples is disclosed.

[0139] According to an aspect of this disclosure, a system (100) for manufacturing a preform (194) of a part of an aircraft (1902) is disclosed, the system comprising:

[0140] Wide-width goods station (110) acquires a sheet (118) made of wide-width goods fiber reinforced material (118-1) and trims the sheet to form a plurality of cladding pieces (116) with boundaries (119).

[0141] A rotating stage (150) aligns the boundary (119) with the boundary (150).

[0142] A shuttle device (170) that holds the laminating piece (116) from the rotary table in a laminating pattern (174); and

[0143] Placement station (181-1) transfers the coating pattern (174) to the coating mandrel (190) and shapes the coating pattern (174) to match the shape (192) of the coating mandrel (190).

[0144] Optionally, the boundaries (119) of the cladding member (116) are complementary to each other.

[0145] Optionally, the placement station (181-1) forms an interleaved splice (177) between the coating component (116) of the coating pattern (174) and the coating component (116) of another coating pattern (174) disposed at the coating mandrel (190).

[0146] Optionally, the layering pattern (174) forms a layer of the wing skin (900).

[0147] Optionally, the wide-width goods station applies a cut to the sheet to produce the cladding (116) exhibiting a common leading edge (195) angle and a common trailing edge (197) angle.

[0148] Optionally, the rotary table (150) rotates the plaster (116) by orienting the leading edge (195) of the plaster (116) to a common angle.

[0149] Optionally, each coating element (116) occupies a zone (902) of the coating pattern (174).

[0150] Optionally, the placement station (181-1) shapes the coating pattern (174) by applying a vacuum to the coating pattern (174).

[0151] Optionally, the cladding pattern (174) forms the shape of the wing skin (900).

[0152] Optionally, the cladding pattern (174) forms the shape of the fuselage skin (1852).

[0153] Optionally, the cladding pattern (174) includes multiple layers (175) made of the wide-width cargo fiber reinforcement material (118-1).

[0154] Optionally, the placement station (181-1) places the coating component (116) in a just-in-time (JIT) manner.

[0155] Optionally, the placement station (181-1) includes a carrier (180) that shapes the coating pattern (174) and conveys the coating pattern (174).

[0156] Optionally, the wide-width goods station (110), the rotary table (150), and the shuttle device (170) are operated repeatedly until the preform (194) is completed at the coating mandrel (190).

[0157] According to an aspect of this disclosure, a part of an aircraft (1902) is disclosed using a system (100) described according to any of the foregoing examples.

[0158] Although specific embodiments have been described herein, the scope of this disclosure is not limited to those specific embodiments. The scope of this disclosure is defined by the appended claims.

Claims

1. A method for manufacturing a preform for a part of an aircraft, the method comprising the steps of: Obtain sheets made of wide-width fiber-reinforced material; Trim the sheet to form a cladding with boundaries; Position the boundary to align; The laminating components are configured according to the laminating pattern to form a layer; Perform a placement operation to transfer the application pattern to the application mandrel; as well as The lamination pattern is shaped to conform to the shape of the lamination mandrel, wherein the trimming step includes: applying a straight cut across the entire width of the sheet at an angle corresponding to either the leading edge angle or the trailing edge angle of the wing skin. The step of trimming the sheet produces the laminate exhibiting a common leading edge angle and a common trailing edge angle, and in the step of setting, the laminates are butted together to form a uniform trailing edge and a uniform leading edge.

2. The method according to claim 1, wherein: The boundaries of the cladding components are complementary to each other.

3. The method according to claim 1 or 2, wherein: The placement operation is performed by forming an interlocking joint between the layering piece of the layering pattern and the layering piece of another layering pattern located at the layering mandrel.

4. The method according to claim 1 or 2, wherein: The step of aligning the boundary includes rotating the dressing by orienting the leading edge of the dressing to a common angle.

5. The method according to claim 1 or 2, wherein: The step of aligning the boundary includes: rotating the dressing by orienting the trailing edge of the dressing to a common angle, and / or Wherein, the sum of the shared leading edge angle and the shared trailing edge angle is not equal to 180°, and / or Each layer occupies a zone of the layer pattern.

6. The method according to claim 1 or 2, wherein: The layering pattern forms the layers of the wing skin, and / or The steps of forming the coating pattern include: applying a vacuum to the coating pattern via a carrier during the placement operation, and / or The step of setting the laminating component to the laminating pattern includes: setting the laminating component to the shape of at least one layer of the wing skin, and / or The step of setting the laminating component to the laminating pattern includes: setting the laminating component to the shape of at least one layer of the fuselage skin, and / or The operations of acquiring, trimming, placing, setting, executing, and shaping are performed via workstations on an assembly line that operate synchronously in a pulsed manner, where each pulse is followed by a pause, and / or The step of forming the plating pattern is performed by a carrier for performing the placement operation, and / or The step of performing the placement operation of the layer pattern is performed in a just-in-time (JIT) manner, and / or the method further includes the following steps: Before performing the placement operation, set an additional layer style on the layer style, and / or The process of trimming, placing, setting, executing, and shaping is repeated until a preform is completed at the coating mandrel.

7. The method according to claim 1 or 2, wherein: The operations of acquiring, trimming, placing, setting, executing, and shaping are performed according to the time of the tick.

8. The method according to claim 7, wherein: The cycle time is synchronized with one or more feeder lines, and / or The cycle time is different for one or more feeder lines.

9. The method according to claim 1 or 2, further comprising the following step: The composite structure is divided into zones; and Manufacturing the coatings for each zone in the said zones, and / or wherein, Each layer occupies a zone of the layer pattern, and / or The steps of transmitting the coating pattern include: operating the pick-and-place PNP station, and / or wherein, The step of setting the cladding component to the cladding pattern includes: setting the cladding component to the shape of a wing skin, and / or wherein, The step of setting the coating component to the coating pattern includes: setting the coating component to the shape of a fuselage skin, and / or The steps of dividing, manufacturing, setting, conveying, and compacting are repeatedly performed until a preform is completed at the coating mandrel, and / or the method further includes the following steps: The edges of the application pattern are positioned on the application mandrel such that the edges stagger relative to other application patterns on the application mandrel, and / or wherein, The additional coating pattern is compacted onto the coating pattern.

10. The method according to claim 9, wherein, The additional coating pattern is staggered with the coating pattern.

11. The method according to claim 9, wherein: The manufacturing, setting, conveying, and compaction steps are performed according to a cycle time.

12. The method according to claim 11, wherein, The cycle time is synchronized with one or more feeder lines, or The cycle time is different for one or more feeder lines.

13. A portion of an aircraft, said portion being assembled by the method according to any one of claims 1 to 12.

14. A system for manufacturing a preform of a part of an aircraft, the system comprising: A wide-width goods processing station, wherein the wide-width goods processing station acquires sheets made of wide-width goods fiber-reinforced material and trims the sheets to form multiple layers with boundaries; A rotary table that positions the boundary for alignment; A shuttle device that holds the plating piece from the rotary table in a plating pattern; as well as A placement station transfers the lamination pattern to a lamination mandrel and shapes the lamination pattern to conform to the shape of the lamination mandrel. The wide-sheet workstation is configured to apply a straight cut across the entire width of the sheet at an angle corresponding to either the leading edge angle or the trailing edge angle of an wing skin. The operation of trimming the sheet produces the laminate exhibiting a common leading edge angle and a common trailing edge angle, and in the operation of shaping the laminate pattern, the laminates are joined together to form a uniform trailing edge and a uniform leading edge.

15. The system according to claim 14, wherein: The boundaries of the cladding components are complementary to each other, and / or The placement station forms an alternating splicing between the coating component of the coating pattern and the coating component of another coating pattern located at the coating mandrel, and / or The layering pattern forms the layers of the wing skin, and / or The rotary table rotates the plaster by orienting the leading edge of the plaster to a common angle, and / or The sum of the shared leading edge angle and the shared trailing edge angle is not equal to 180°, and / or Each layer occupies a zone of the layer pattern, and / or The placement station shapes the coating pattern by applying a vacuum to the coating pattern, and / or The cladding pattern forms the shape of the wing skin, and / or The layering pattern forms the shape of the fuselage skin, and / or The cladding pattern includes multiple layers made of the wide-width cargo fiber reinforcement material, and / or The placement station places the coating component using a just-in-time (JIT) method, and / or The placement station includes a carrier that shapes and transports the application pattern, and / or The wide-width goods station, rotary table, and shuttle device are repeatedly operated until the preform is completed at the lamination mandrel, and / or The wide-width workstation applies a cut to the sheet material to produce the laminate exhibiting a common leading edge angle and a common trailing edge angle.

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