Method for simultaneous manufacturing of gasketless components

Abstract

The present application relates to a method for simultaneous manufacturing of shim-free components. The method is intended to reduce and / or eliminate the need for shims in manufacturing components, such as in the manufacture of aircraft wings. An exemplary method includes: predicting a set of predicted manufacturing dimensions within a predetermined allowable tolerance range of a first component; manufacturing the first component; scanning the first component to determine a set of actual manufacturing dimensions of the first component; and completing and starting manufacturing of a second component at least before scanning the first component. The second component can be manufactured based on the set of predicted manufacturing dimensions of the first component. Once the scanning of the first component is completed, the set of predicted manufacturing dimensions can be compared with the set of actual manufacturing dimensions to check whether there are any undesirable deviations between the predicted manufacturing dimensions and the actual manufacturing dimensions. Repairs and local rescans can be performed in areas of undesirable deviations, which can simplify the manufacturing process.

Description

Technical Field

[0001] The present disclosure generally relates to methods of simultaneous manufacturing of shimless assemblies. Background Art

[0002] During the manufacturing process, multiple structural components are often manufactured separately and then assembled together. For example, in an aerospace or aviation application, such as in the manufacture of an aircraft wing, dozens or even hundreds (or more) of structural components may be assembled together to form a structure. Gaps between such structural components may be caused by manufacturing tolerances (expected but unplanned deviations) of the components and / or unique challenges associated with certain materials. For example, when manufacturing a composite component, geometric variations in the final component may be caused by variations in fiber diameter and / or variations in resin volume, which may be accumulated via multiple layers of material that are stacked to form the composite component.

[0003] One example where such a problem arises is the assembly of skin panels to rib feet to form the wing box of an aircraft. Figure 1 An example of a wing box structure for an aircraft wing 10 is shown. Typically, the aircraft wing 10 includes a ladder-like structure formed by a plurality of ribs 12 spaced apart between one or more longer spars 14. The ribs 12 define the overall shape of the aircraft wing 10, wherein a skin panel 16 (partially shown in phantom, also referred to as a wing skin) is attached to the ribs 12, conforming to the shape of the ribs 12. In conventional processes, the skin panel 16 is made into a desired shape and then placed in place to engage the ribs 12 and the spars 14. The underlying rib / spar ladder structure may have, for example, outwardly protruding rib feet, on which portions of the inner surface of the skin panel 16 are intended to rest, so that fasteners may be inserted through the skin panel 16 and the rib feet to secure the skin panel to the ladder-like structure. The engineering tolerances of the skin panel 16 are typically greater than the acceptable error for assembly between the skin panel 16 and the ladder-like structure, which may result in interface gaps between the skin panel 16 and the ladder-like structure in certain areas. For example, when the skin panel is placed in an unstressed state against the rib feet, it is often found that while some of the rib feet are in contact with the skin panel, other rib feet are spaced apart from the skin panel. The skin panel can only deform slightly to safely accommodate these gaps without generating stresses stored within the skin panel or affecting the shape and aerodynamics of the skin panel. However, enforcing extreme tolerances increases costs and / or may simply not be feasible. Therefore, manufacturers typically rely on filling the gaps by applying liquid or solid gaskets between the ribs 12 and the skin panel 16 when needed (e.g., at the location of the gap).

[0004] To produce the required shims, the skin panel and the substructure of the wing box are usually placed in place adjacent to each other so that the gap between the rib foot and the inner surface of the skin panel can be measured. Shims are then manufactured to fill the gaps, with the shape and size of each shim being selected according to the shape and size of the respective gap to be filled. In this method, the final assembly of the skin panel and the substructure of the wing box is postponed until after the shims are manufactured, at which time the skin panel and the substructure will be placed in place adjacent to each other again. Such added steps and delays increase manufacturing costs and reduce efficiency. In addition, placing these shims is often a time-consuming and expensive process. In some cases, the composite structure may have to be assembled and disassembled multiple times to measure the shim gaps and drill and clean the holes. Care must also be taken to ensure that shims that have been manufactured to a specific size and shape to fill a specific gap are used to fill the correct gap and are not inadvertently used to fill other gaps.

[0005] In some manufacturing processes, even more delays are introduced at other stages. For example, efforts have been made to reduce the use of shims, such as by manufacturing the skin panels and then custom manufacturing the spars and spar caps after the as-built dimensions of the skin panels are known. This approach can create an inventory buildup of wing skin panels while the spars and spar caps are manufactured. Wing skin panels are large structures that must be properly held and supported while the spars and spar caps are manufactured to prevent damage to the skin panels, which can take up valuable warehouse space for a long time. Even with such attempts, shims may still be required in some places. Summary of the invention

[0006] The presently disclosed method is intended to reduce and / or eliminate the need for shims in manufacturing assemblies, such as in the manufacture of aircraft wings. An exemplary method of manufacturing an assembly of at least a first component and a second component generally includes: predicting a set of predicted manufacturing dimensions within a predetermined allowable tolerance range of the first component; manufacturing the first component; scanning the first component to determine a set of actual manufacturing dimensions of the first component; and starting manufacturing the second component at least before scanning of the first component is completed. The set of predicted manufacturing dimensions may be predicted using a historical data set from a plurality of previously built components, the data set including dimensions of the previous components and data from a plurality of factors. The second component may be manufactured based on the set of predicted manufacturing dimensions of the first component, such that the second component is configured to engage the predicted first component. However, such a method may save valuable manufacturing time by having the second component at least partially or completely completed before the first component has been fully scanned (and therefore before the actual as-built dimensions of the respective first components are known). Once the scanning of the first component is completed (and the as-built dimensions of the first component are determined), the set of predicted manufacturing dimensions may be compared with the set of actual manufacturing dimensions to check whether there are any undesirable deviations between the set of predicted manufacturing dimensions and the set of actual manufacturing dimensions. Such a method can achieve automated manufacturing of the second component (e.g., without the need to customize the second component during the initial manufacturing stage), and thus can simplify the entire assembly process and only modify (e.g., customize) the second component within any undesirable deviation areas detected between the built first component and the second component.

[0007] Also disclosed is a component scan model that can be configured to create a model of an interface region between a first mating structure and a second mating structure. Such a model can include an original large scan base layer of the entire interface region at the time of initial manufacture, and one or more small scan layers stitched into the large scan base layer, wherein the small scan layers correspond to local portions of the interface region that are modified after initial manufacture. BRIEF DESCRIPTION OF THE DRAWINGS

[0008] Figure 1 is a perspective view of a prior art example of a structure for an aircraft wing.

[0009] Figure 2 is a schematic flow chart illustrating a representative method of manufacturing an assembly according to the present disclosure.

[0010] Figure 3 is an example of a portion of an assembly that may be manufactured according to the method of the present disclosure, shown partially exploded. DETAILED DESCRIPTION

[0011] Figure 2 Schematically provided is a flow chart representing an illustrative, non-exclusive example of a method 100 according to the present disclosure. Figure 2In the embodiment of the present invention, some steps are shown in dashed boxes, indicating that these steps may be optional or may correspond to optional versions of the method according to the present disclosure. That is, it is not required that all methods according to the present disclosure include the steps shown in solid boxes. It can be understood from the discussion herein that Figure 2 The methods and steps shown in are not limiting, and other methods and steps are also within the scope of the present disclosure, including methods having more or less than the number of steps shown.

[0012] The method 100 described herein is generally applicable to manufacturing an assembly of two or more components, such as assembling a first component and a second component. Of course, such a method is generally applicable to manufacturing an assembly of multiple components, but for the sake of clarity, it will first be described in terms of assembling a first component (e.g., a wing skin panel for an aircraft wing) and a second component (e.g., a rib for an aircraft wing). Figure 3 A partial exploded portion of an aircraft wing assembly 20 is shown for illustration purposes only. Figure 2 Reference will be made to the discussion of the method 100 of . The wing assembly 20 includes a skin panel 22 (which may also be referred to as a wing panel skin, a wing skin panel, a panel skin, a wing skin, or a wing panel), a first rib 24, and first and second spar caps 26a, 26b. As will be appreciated by those skilled in the art, the wing assembly 20, when completed, includes a plurality of ribs spaced along the length of the skin panel 22 in addition to one or more other skin panels to complete the structure. The spar caps 26a, 26b are typically secured to the skin panel 22 along a longitudinal interface region 28 extending longitudinally of the skin panel 22, while the first rib 24 is secured to the skin panel 22 along a transverse interface region 30 extending transversely across the skin panel 22, as shown in FIG. Figure 3 Any additional ribs secured to the skin panel 22 may be secured to other corresponding lateral interface regions spaced apart from the lateral interface region 30. In examples involving assembly of different structures other than the exemplary aircraft wing, the first component and the second component will typically have interface regions that differ in number, location, spacing, shape, size, and / or orientation, but such variations do not depart from the scope of the present disclosure.

[0013] The disclosed method 100 includes predicting a set of predicted manufacturing dimensions of a first component, such as a set of predicted manufacturing dimensions of a skin panel 22, at 102. Although the manufactured component has a designed or expected set of dimensions, it should be understood that the actual built component may be slightly different from the designed dimensions. For example, although the manufacturing process may intend to produce a plurality of fasteners each having a length of 1.00 centimeters (cm), the length of the fasteners actually produced may vary between, for example, 0.95 cm and 1.05 cm. For components with complex contours and shapes, this variation also increases complexity. Therefore, the set of predicted manufacturing dimensions may be a predicted range of manufacturing dimensions. In other words, a predetermined and / or acceptable allowable tolerance range (i.e., an acceptable specification, variation, and / or size range) may be predicted at 102 based on a historical data set from a plurality of previously built components. For example, in the case of a skin panel 22, the predicted manufacturing dimensions may be predicted at 102 based on a historical data set from a plurality of previously built skin panels (as opposed to simply looking at the designed dimensions of the skin panel). Predicting the predicted manufacturing dimensions at 102 typically includes characterizing the interface where the components will be joined to each other.

[0014] The predicted manufacturing dimensions are typically predicted at 102 based on a number of independent factors, rather than simply being an average of the historical dimensions of the components that have been built. For example, the prediction may be based on the supplier of the material used to manufacture the component, the specific batch of material used, the atmospheric and / or local environmental conditions in which the component is manufactured, the type of bagging material or other tooling material used in the layup process, the type of prepreg material used, the end effector (or its settings) used in the manufacturing process, the feed rate of the tape laying head used to lay the fiber tape, and / or the temperature and / or pressure used in the autoclave when manufacturing the component. Such data associated with manufacturing the component is typically recorded, and the currently disclosed method may utilize any available information to predict the dimensions of the future component. Determining the predicted manufacturing dimensions at 102 may include statistical process control, and / or may utilize analysis and / or machine learning in some methods 100. Such a predictive model may take many inputs (e.g., any or all of the factors listed above) in order to predict the dimensions of the manufactured component. For example, the predictive model may predict that the dimensions of a component manufactured using the material of a first supplier are different from the dimensions of the same component manufactured using the material of a second supplier. Advantageously, the currently disclosed method 100 may allow for cheaper manufacturing of the first component. For example, in some methods 100 , predicting the predicted dimension at 102 may result in a sufficiently accurate prediction even when the first component uses a less accurate (and cheaper) manufacturing technique.

[0015] Determining the predicted dimensions at 102 may be performed before, during, and / or after manufacturing the first component (e.g., the skin panel 22) at 104. Manufacturing the first component at 104 may include steps such as trimming one or more layers of the first component, drilling one or more holes in one or more layers of the first component, and / or laying up a plurality of fiber reinforced polymer layers. The first component may be scanned at 106 to measure and determine a set of actual manufacturing dimensions of the built first component. Typically, scanning the first component at 106 includes performing a global scan of the entire component, although in some examples, only a portion of the first component may be scanned at 106. Scanning the first component at 106 may be performed via, for example, laser scanning, optical scanning, photography, metrology scanning, and / or any other suitable technology. In some examples, scanning the first component at 106 may include taking multiple photos or images of the first component (e.g., using a large camera platform, and / or hanging and / or moving the camera relative to the first component), with reference markers located on the first component so that the reference markers can be used as reference points for combining the images, stitching the images together to form a composite scan of the first component. Scanning the first component at 106 generally includes at least scanning any interface regions where the second component will be joined to the first component. For example, in the case of a skin panel 22, scanning the first component at 106 may include scanning one or more longitudinal interface regions 28 (e.g., scanning a spar interface) at 108, and / or scanning one or more transverse interface regions 30 (e.g., at scanning rib locations) at 110. The scanning of the first component at 106 may also be performed for quality control, documentation, learning, and / or future analysis purposes.

[0016] The method 100 also includes manufacturing one or more second components (and / or one or more additional components) for the assembly at 112, which may be initiated at least before the scanning of the first component at 106 is completed. For example, the manufacturing of the second component at 112 may be at least partially performed or fully completed during the scanning of the first component at 106 (e.g., before the scanning of the first component is completed). Although not intuitive, such sequencing may improve manufacturing time and efficiency. In the current example, manufacturing the second component at 112 may include manufacturing one or more spars and / or spar caps (e.g., spar caps 26a, 26b) at 114 and / or manufacturing one or more ribs (e.g., first rib 24) at 116. The second component may be manufactured at 112 such that the second component is configured to engage the first component based on the predicted manufactured dimensions of the first component. In other words, the second component may be manufactured at 112 before the complete as-built dimensions of the first component are known. This may allow for automation of manufacturing of the second component as opposed to waiting to determine the as-built dimensions of the first component before starting custom manufacturing of the second component.

[0017] After the scanning of the first component is completed at 106, the actual manufactured dimensions of the built first component (determined by scanning the first component at 106) are compared to the predicted manufactured dimensions of the first component at 118. By such a comparison, any undesirable deviations between the actual manufactured dimensions and the predicted manufactured dimensions can be detected. The criteria for undesirable deviations can be predetermined based on the desired tolerances and allowances. For example, undesirable deviations can be determined in areas where each actual built dimension is greater than the largest dimension within the predicted dimension range, or less than the smallest dimension within the predicted dimension range. In addition, the actual manufactured dimensions of the first component can be added to the historical data set at 120 for use in predicting the dimensions of the first component that is subsequently manufactured (e.g., for future data analysis of future components).

[0018] The first component and the second component (and any other parts) are joined together, or assembled at 122, to form an assembly. For example, the skin panel 22, the first rib 24, and the spar caps 26a, 26b ( Figure 3 ) are assembled with additional components to form the aircraft wing assembly 20. In some methods 100, the components may be mated to one another (e.g., brought together to join one another at an interface region) without requiring metrology or scanning of a second component (e.g., a ladder structure of ribs and spars) prior to joining the second component to a first component (e.g., a skin panel).

[0019] In the method 100 including the third component (and / or additional second components, and / or additional other components), the third component may also be manufactured based on a set of predicted manufacturing dimensions of the first component to join the first component. For example, before the skin panel 22 is fully scanned, the second component (e.g., the first rib 24) and the third component (e.g., the first spar cap 26a) may be manufactured separately according to the predicted dimensions of the skin panel 22. In some methods 100, the manufacturing of the components for assembly can be said to be separate compared to conventional techniques. For example, the spar caps 26a, 26b may be initially manufactured to the same data set each time without relying on the as-built dimensions of the ribs and skin panels when the spar caps are manufactured.

[0020] In some methods 100, comparing the set of predicted dimensions to the set of actual dimensions at 118 may result in one or more unsatisfactory deviations being detected at 124. In such a method 100, repairs may be performed at 126 to accommodate or eliminate the unsatisfactory deviations detected at 124. For example, the first component, the second component, and / or the third component may be modified so that the actual dimensions are within the predicted range of dimensions so that the components may be joined and assembled without the need for shimming to be placed between the components (e.g., so that the fit between the first component and the second component is within a predetermined tolerance). In the example of an aircraft wing assembly, unsatisfactory deviations may be detected where the as-built dimensions of the components would result in an unacceptable interface gap between the skin panel 22 and the first rib 24 and / or between the skin panel 22 and the spar caps 26a and / or 26b, which would otherwise require shimming to be placed at the unsatisfactory interface. However, in the presently disclosed method, rather than shimming for such unsatisfactory deviations, the components may be locally repaired to prevent the need for shimming. For example, the skin panel 22 may be trimmed, "faced," and / or machined down (e.g., material removed), material may be removed from the first rib 24, and / or material may be removed from the spar caps 26a and / or 26b so that the components may be assembled and joined to one another within predetermined tolerances. In some cases, repairing the components at 126 may include adding material to the first component, the second component, and / or any other components, although as used herein, the phrase "adding material" does not include inserting shims into the assembly. For example, sacrificial material may be added to the component in the deviation area that does not meet the requirements and then machined.

[0021] Repairing the component at 126 may include determining how much material to add to or remove from one or more contact surfaces. For example, repairing the component at 126 may include adding material from or removing material from the skin panel 22 at one or more different locations along a given transverse interface region 30 and / or at one or more different locations along a given longitudinal interface region 28. Additionally or alternatively, repairing the component at 126 may include adding material to or removing material from a contact surface of the first rib 24 configured to engage the transverse interface region 30 and / or adding material to or removing material from a contact surface of the spar caps 26a and / or 26b configured to engage the longitudinal interface region 28. In other words, repairing the component at 126 may include modifying one or more corresponding local regions of the first component, one or more corresponding local regions of the second component, and / or one or more corresponding local regions of the third component, wherein the modified local regions correspond to regions that engage one or more other portions of the detected non-compliant deviation region. Because the second and third components have been manufactured according to predicted dimensions (as opposed to designed or expected dimensions) in method 100 , machining the components for the repair at 126 is generally faster and more efficient than creating shims in conventional methods.

[0022] After the component is repaired at 126, one or more local areas may be re-scanned at 128 in the area of ​​unsatisfactory deviation detected at 124 (e.g., the local area where the modification and / or repair was performed at 126). The re-scanning at 128 may include re-scanning the first component at the interface location corresponding to the unsatisfactory deviation and / or re-scanning the assembly of the component in the area of ​​unsatisfactory deviation. Advantageously, the repaired portion may be locally re-scanned at 128 without placing the component in a large rack, which may save costs in some methods 100. As a result of the local scanning at 128, if unsatisfactory deviations are still detected at 124, additional repairs may be performed at 126, and repeated until the final component assembly is within tolerance. At 130, the data (e.g., dimensions) from the local scan at 128 may be digitally inserted (which may be referred to herein as digitally stitched) into the set of actual manufacturing dimensions so that the set of actual manufacturing dimensions is accurate for the final built component, and then added to the historical data set at 120. For example, the reference coordinates of the first component may be used during the rescan at 128 and / or the scan at 106 to facilitate inserting data from the rescan at 128 into the larger scan of the first component.

[0023] Thus, the method 100 may include creating a component scan model as part of the historical data set, wherein the component scan model includes a scan of an interface region between a first mating structure and a second mating structure, such as Figure 3 The longitudinal and / or transverse interface regions between the skin panels and ribs and / or spar of an aircraft wing assembly. In one example, the first mating structure may be a spar interface (e.g., longitudinal interface region 28) of a skin panel for an aircraft wing, and the second mating structure may be a spar cap (e.g., spar cap 26a, 26b) configured to engage the skin panel. Additionally or alternatively, the first mating structure may be a rib position (e.g., transverse interface region 30) of a skin panel for an aircraft wing, and the second mating structure may be a rib (e.g., first rib 24) configured to engage the skin panel. The component scan model may include an original large scan base layer of the entire interface region (e.g., the entire longitudinal interface region 28 and / or the entire transverse interface region 30) at the time of initial manufacturing, and one or more small scan layers sewn into the large scan base layer. The small scan layer corresponds to a local portion of one or more interface regions that are modified after initial manufacturing, such as via a repair or modification performed at 126 in response to a detected non-compliant deviation.

[0024] Such a method 100 according to the present disclosure may provide numerous advantages in manufacturing, particularly in manufacturing large assemblies, such as aircraft wings. By starting manufacturing the second component at 112 before completing scanning of the first component at 106, the factory's inventory may be reduced by reducing the amount of time that the skin panels need to be stored prior to assembly with other components. This reduction in required storage space, as well as the reduction in manufacturing time due to not having to form, shape, and place multiple shims in the assembly, may provide cost savings for a given manufacturing process. Some methods 100 may also result in a reduction in the amount of rework required and / or an increase in overall component quality in the finished component, as compared to prior art techniques involving the placement of shims. The method 100 may also result in increased manufacturing speed (e.g., throughput) and / or reduced complexity in manufacturing logistics by eliminating the need to place shims in the assembly.

[0025] Illustrative, non-exclusive examples of inventive subject matter according to the present disclosure are described in the following enumerated paragraphs:

[0026] A1. A method of manufacturing an assembly comprising a first component and a second component, the method comprising:

[0027] predicting a set of predicted manufacturing dimensions within a predetermined allowable tolerance range for said first component using a historical data set from a plurality of previously built components;

[0028] manufacturing the first component;

[0029] scanning the first component to determine a set of actual manufactured dimensions of the first component;

[0030] manufacturing the second component, wherein the second component is manufactured based on the set of predicted manufacturing dimensions of the first component such that the second component is configured to engage the first component, wherein manufacturing the second component is at least partially performed before scanning of the first component is completed; and

[0031] The set of predicted manufacturing dimensions is compared with the set of actual manufacturing dimensions to check whether there are any undesirable deviations between the set of predicted manufacturing dimensions and the set of actual manufacturing dimensions.

[0032] A1.1. The method of paragraph A1, wherein manufacturing the second component is completed before scanning the first component is completed.

[0033] A1.2. The method of paragraph A1 or A1.1, wherein scanning the first component includes performing a global scan of the entire first component.

[0034] A1.3. The method of any of paragraphs A1-A1.2, further comprising adding the actual manufacturing dimensions to the historical data set.

[0035] A2. The method of any of paragraphs A1-A1.3, further comprising bonding the second component to the first component to form the assembly.

[0036] A3. The method of any of paragraphs A1-A2, wherein the first component comprises a skin panel for an aircraft wing.

[0037] A4. The method of any of paragraphs A1-A3, wherein the second component comprises a rib for a wing of the aircraft.

[0038] A5. The method of any of paragraphs A1-A4, wherein the second component comprises a spar cap for a wing of the aircraft.

[0039] A6. The method of any of paragraphs A1-A5, wherein the assembly includes a third component, wherein the method further comprises:

[0040] manufacturing the third component, wherein the third component is manufactured to be configured to engage with the first component based on the set of predicted manufacturing dimensions of the first component, wherein manufacturing the third component is at least partially performed before scanning of the first component is completed; and

[0041] The third component is joined to the first component to form the assembly.

[0042] A6.1. The method of paragraph A6, wherein manufacturing the third component is completed before scanning the first component is completed.

[0043] A7. The method of paragraph A6 or A6.1, wherein the first component comprises a skin panel for an aircraft wing, wherein the second component comprises a rib for an aircraft wing, and wherein the third component comprises a spar cap for an aircraft wing.

[0044] A8. The method according to any of paragraphs A1-A7, further comprising:

[0045] detecting one or more undesirable deviations between the set of predicted manufacturing dimensions and the set of actual manufacturing dimensions; and

[0046] The second component and / or the third component are modified to accommodate the one or more undesirable deviations such that the second component and / or the third component engages the first component within a predetermined tolerance when assembled.

[0047] A9. The method of paragraph A8, wherein modifying includes one or more selected from the group consisting of: removing material from the second component, removing material from the third component, adding material to the second component, and adding material to the third component.

[0048] A9.1. The method of paragraph A8 or A9, wherein the modifying includes determining how much material to remove and / or add from one or more corresponding contact surfaces of the second component and / or one or more corresponding contact surfaces of the third component.

[0049] A10. A method according to any of paragraphs A8-A9.1, wherein the modification includes modifying one or more corresponding local areas of the second component and / or one or more corresponding local areas of the third component, wherein the local areas correspond to areas that are engaged with the first component in the area where the undesirable deviation is detected.

[0050] A.11. The method according to any of paragraphs A8-A10 also includes: local scanning of one or more local areas corresponding to the detected non-compliant deviations.

[0051] A12. The method of paragraph A11, wherein a local scan of the one or more local areas is performed after modifying the second component and / or the third component.

[0052] A13. The method of paragraph A11 or A12 further includes digitally inserting data from the local scan of the one or more local areas into the set of actual manufacturing dimensions.

[0053] A14. A method according to any of paragraphs A1 to A13, wherein the prediction is affected by multiple independent factors.

[0054] A14.1. A method as described in paragraph A14, wherein the independent factors include the supplier of the material, the specific batch of material, atmospheric and / or local environmental conditions, the type of bagging material or other tooling material used in the layup process, the type of prepreg material used, the end effector (or its configuration) used in the manufacturing process, and / or the temperature and / or pressure used in the autoclave when manufacturing the first part.

[0055] A15. A method as described in any of paragraphs A1-A14.1, wherein the prediction includes statistical process control.

[0056] A16. The method of any of paragraphs A1-A15, wherein manufacturing the second component is automated without requiring custom machining of the second component to engage with a specific configuration of the first component.

[0057] A17. The method of any of paragraphs A1-A16, wherein manufacturing the first component includes:

[0058] laying down multiple fiber-reinforced polymer layers;

[0059] Pruning layers; and

[0060] Drill multiple holes in the layer.

[0061] A18. A method according to any of paragraphs A1-A17, wherein the first component includes a skin panel for an aircraft wing, wherein the skin panel includes one or more spar interfaces for engaging one or more spars, wherein the skin panel includes one or more rib locations for engaging one or more ribs, and wherein the scanning includes scanning the spar interfaces and scanning the rib locations.

[0062] B1. A component scanning model, comprising:

[0063] A scan of an interface region between a first mating structure and a second mating structure, wherein the scan comprises:

[0064] The original large scan base layer across the interface area at the time of initial fabrication; and

[0065] One or more small scan layers are stitched into the large scan base layer, wherein the small scan layers correspond to localized portions of the interface region modified after initial fabrication.

[0066] B2. The component scanning model of paragraph B1, wherein the first mating structure comprises a spar interface for a skin panel of an aircraft wing, and wherein the second mating structure comprises a spar configured to engage the skin panel.

[0067] B3. The component scanning model of paragraph B1, wherein the first mating structure comprises a rib location for a skin panel of an aircraft wing, and wherein the second mating structure comprises a rib configured to engage the skin panel.

[0068] D1. Use the method described in any of paragraphs A1-A18 to join multiple parts together.

[0069] D2. Use the method described in any of paragraphs A1-A18 to manufacture an aircraft wing.

[0070] D3. Connect one or more ribs to the skin panel using the method described in any of paragraphs A1-A18.

[0071] D4. Connect one or more spars and / or spar caps to the skin panel using the method described in any of paragraphs A1-A18.

[0072] As used herein, the terms "selective" and "selectively" when used in connection with an action, motion, configuration, or other activity that modifies one or more components or characteristics of a device, mean that the particular action, motion, configuration, or other activity is the direct or indirect result of a user operating on an aspect or one or more components of the device.

[0073] As used herein, the terms "suitable for" and "configured to" refer to an element, component, or other subject matter that is designed and / or intended to perform a given function. Therefore, the use of the terms "suitable for" and "configured to" should not be interpreted as meaning that a given element, component, or other subject matter is merely "capable of" performing a given function, but rather that the element, component, and / or other subject matter is specifically selected, created, implemented, utilized, programmed, and / or designed to perform the function. It is also within the scope of the present disclosure that an element, component, and / or other narrated subject matter that is described as being suitable for performing a particular function may additionally or alternatively be described as being configured to perform that function, and vice versa. Similarly, a subject matter that is described as being configured to perform a particular function may additionally or alternatively be described as being operable to perform that function.

[0074] As used herein, the phrase "at least one" with respect to a list of one or more entities should be understood to refer to at least one entity selected from any one or more entities in the list of entities, but does not necessarily include at least one of each entity explicitly listed in the list of entities, and does not exclude any combination of entities in the list of entities. This definition also allows for the optional presence of entities in addition to the entities specifically identified within the list of entities to which the phrase "at least one" refers, whether related or unrelated to those specifically identified entities. Thus, as a non-limiting example, "at least one of A and B" (or equivalently, "at least one of A or B", or equivalently "at least one of A and / or B") may refer to at least one A, optionally including more than one, with no B (and optionally including entities other than B) in one embodiment; at least one B, optionally including more than one, with no A (and optionally including entities other than A) in another embodiment; and at least one A, optionally including more than one, and at least one B, optionally including more than one (and optionally including other entities) in yet another embodiment. In other words, the phrases "at least one", "one or more", and "and / or" are open-ended expressions that are conjunctive and disjunctive in operation. For example, each expression "at least one of A, B, and C", "at least one of A, B or C", "one or more of A, B and C", and "A, B and / or C" may mean A alone, B alone, C alone, A and B together, A and C together, B and C together, A, B and C together, and optionally any of the above in combination with at least one other entity.

[0075] Not all devices and methods according to the present disclosure require the steps of the various disclosed elements and methods of the devices disclosed herein, and the present disclosure includes all novel and non-obvious combinations and sub-combinations of the various elements and steps disclosed herein. In addition, one or more of the various elements and steps disclosed herein may define an independent inventive subject matter, which is independent of and different from the entirety of the disclosed device or method. Therefore, such an inventive subject matter is not required to be associated with the specific devices and methods explicitly disclosed herein, and such an inventive subject matter may find utility in devices and / or methods not explicitly disclosed herein.

[0076] As used herein, the phrase "for example", the phrase "as an example" and / or simply the term "example", when used with reference to one or more components, features, details, structures, implementations, and methods according to the present disclosure, is intended to convey that the described components, features, details, structures, implementations, and / or methods are illustrative, non-exclusive examples of components, features, details, structures, implementations, and / or methods according to the present disclosure. Thus, the described components, features, details, structures, implementations, and / or methods are not intended to be limiting, requiring, or exclusive / exhaustive; and other components, features, details, structures, implementations, and / or methods, including components, features, details, structures, implementations, and / or methods that are similar and / or equivalent in structure and / or function, also fall within the scope of the present disclosure.

Claims

1. A method (100) of manufacturing an assembly (20) comprising a first component and a second component, the method (100) comprising: Step 1 (102), using a historical data set from a plurality of previously built parts, predicting a set of predicted manufacturing dimensions within a predetermined allowable tolerance range of the first part; Step 2 (104), manufacturing the first component; Step three (106), scanning the first component to determine a set of actual manufacturing dimensions of the first component; Step four (112), manufacturing the second component, wherein the second component is manufactured based on the set of predicted manufacturing dimensions of the first component such that the second component is configured to engage the first component, wherein step four (112) of manufacturing the second component is at least partially performed before step three (106) of scanning the first component is completed; and Step five (118), comparing the set of predicted manufacturing dimensions with the set of actual manufacturing dimensions to check whether there is any deviation that does not meet the requirements between the set of predicted manufacturing dimensions and the set of actual manufacturing dimensions.

2. The method (100) according to claim 1, wherein: The fourth step (112) of manufacturing the second component is completed before the third step (106) of scanning the first component is completed.

3. The method (100) according to claim 1 or 2, wherein: The third step (106) of scanning the first component includes performing a global scan on the entire first component.

4. The method (100) according to claim 1 or 2, further comprising step six (120), adding the set of actual manufacturing dimensions to the historical data set.

5. The method (100) according to claim 1 or 2, further comprising a step seven (122) of bonding the second component to the first component to form the assembly (20).

6. The method (100) according to claim 5, wherein: The first component comprises a skin panel (22) for an aircraft wing, and wherein the second component comprises a rib (24) for the aircraft wing.

7. The method (100) according to claim 6, wherein: The assembly (20) comprises a third component, the third component comprising a spar cap for the aircraft wing, wherein the method (100) further comprises: manufacturing the third component, wherein the third component is manufactured such that the third component is configured to engage with the first component based on the set of predicted manufacturing dimensions of the first component, wherein manufacturing the third component is at least partially performed before step three (106) of scanning the first component is completed; and The third component is joined to the first component to form the assembly (20).

8. The method (100) according to claim 1 or 2, further comprising: Step eight (124), detecting one or more non-compliant deviations between the set of predicted manufacturing dimensions and the set of actual manufacturing dimensions; as well as Step nine (126), modifying the second component to accommodate the one or more undesirable deviations so that the second component engages the first component within a predetermined tolerance when assembled.

9. The method (100) according to claim 8, wherein: The modifying step nine (126) includes one or more selected from the group consisting of: removing material from the second component and adding material to the second component.

10. The method (100) according to claim 9, wherein: The step nine of modifying (126) includes modifying one or more corresponding local regions of the second component that are regions that engage the first component in the region of the detected undesirable deviation.

11. The method (100) according to claim 8, further comprising step ten (128), locally scanning one or more local areas corresponding to the detected non-compliant deviations.

12. The method (100) according to claim 11, wherein: After step nine (126) of modifying the second component, a local scan of the one or more local areas is performed.

13. The method (100) of claim 12, further comprising a step eleven (130) of digitally inserting data from step ten (128) of the local scan of the one or more local areas into the set of actual manufacturing dimensions.

14. The method (100) according to claim 8, wherein: The prediction step one (102) is influenced by a number of independent factors.

15. The method (100) according to claim 1 or 2, wherein: Step one (102) of prediction includes statistical process control.

16. The method (100) according to claim 1 or 2, wherein: Step four (112) of manufacturing the second component is automated without requiring custom machining of the second component to engage with a specific configuration of the first component.

17. The method of claim 1, further comprising creating a component scan model, comprising: A scan of an interface region between a first mating structure and a second mating structure, wherein the scan comprises: The original large scan base layer across the interface area at the time of initial fabrication; and One or more small scan layers are stitched into the large scan base layer, wherein the small scan layers correspond to localized portions of the interface region modified after initial fabrication.

18. The method according to claim 17, wherein: The first mating structure comprises a spar interface for a skin panel (22) of an aircraft wing, and wherein the second mating structure comprises a spar configured to engage the skin panel (22).

19. The method according to claim 17 or 18, wherein: The first mating structure comprises a rib location for a skin panel (22) of an aircraft wing, and wherein the second mating structure comprises a rib configured to engage the skin panel (22).

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