How to synchronize the manufacturing of shimless assemblies.

Abstract

To provide a method that aims to reduce and / or eliminate the need for shims in manufacturing assemblies, such as in manufacturing of aircraft wings.SOLUTION: An exemplary method comprises the steps of: predicting a set of predicted manufacturing dimensions within a range of predetermined allowances for a first part; manufacturing the first part; scanning the first part to determine a set of actual manufacturing dimensions for the first part; and at least beginning manufacturing a second part before the scanning of the first part is completed. The second part may be manufactured based on the set of predicted manufacturing dimensions for the first part. Once the scan of the first part is completed, the set of predicted manufacturing dimensions may be compared to the set of actual manufacturing dimensions to check for any non-compliant deviances between the predicted and actual manufacturing dimensions. Repairs and local re-scans may be performed in the areas of the non-compliant deviances, which may streamline manufacturing.SELECTED DRAWING: Figure 2

Description

[Technical Field]

[0001] FIELD OF THE DISCLOSURE The present disclosure relates generally to methods for synchronizing the manufacture of shimless assemblies. [Background technology]

[0002] Manufacturing processes often involve the assembly of multiple structural components that are manufactured separately. For example, in aerospace and aviation applications, such as the manufacture of aircraft wings, dozens or even hundreds of structural components (or more) may be assembled to form a structure. Gaps may occur between such structural components due to component manufacturing tolerances (variations that are expected but cannot be planned for) and / or unique issues associated with certain materials. For example, when manufacturing a composite part, geometric variations in the finished part may occur due to variations in fiber diameter and / or variations in resin volume that may accumulate through the multiple layers of material that are laid up to form the composite part.

[0003] An example of where such a problem occurs is in the assembly of skin panels to rib feet that form an aircraft wing box. FIG. 1 shows an example wing box structure for an aircraft wing 10. Generally, the aircraft wing 10 includes a ladder-like structure formed by a plurality of spaced ribs 12 between one or more elongated spars 14. Skin panels 16 (partially shown in dashed lines and also referred to as wing skins) are attached to and conform to the shape of the ribs 12, which generally define the overall shape of the aircraft wing 10. In a conventional process, the skin panels 16 are manufactured in the desired shape and then positioned to engage the ribs 12 and spars 14. The underlying rib / spar ladder structure may, for example, have outwardly protruding rib feet against which respective portions of the inner surface of the skin panels 16 rest so that fasteners can be inserted into the skin panels 16 and rib feet to secure the skin panels to the ladder structure. Typically, the fabrication tolerances for the skin panel 16 are greater than the acceptable tolerance for the fit between the skin panel 16 and the ladder structure, which can result in interface gaps between the skin panel 16 and the ladder structure in some areas. For example, it is generally found that when the skin panel is positioned against the rib feet in an unstressed state, some rib feet will abut the skin panel while others will be spaced apart. Only a small amount of deformation of the skin panel can successfully accommodate such gaps without causing stress buildup in the skin panel or affecting the shape and aerodynamics of the skin panel. However, forcing extreme tolerances can increase costs and / or be completely impractical. Therefore, manufacturers typically rely on filling the gaps by applying liquid or solid shims between the ribs 12 and the skin panel 16 at the required locations (e.g., at the gap locations).

[0004] To fabricate the necessary shims, the skin panel and the underlying wing box structure are typically positioned side-by-side so that the gap between the rib foot and the inner surface of the skin panel can be measured. The gap is then filled with the shim. The shape and size of each shim are selected according to the shape and size of the respective gap to be filled. This approach postpones the final assembly of the skin panel and the underlying wing box structure until after the shim is fabricated, at which point the skin panel and the underlying structure are again positioned side-by-side. This additional step and delay increases manufacturing costs and reduces efficiency. Additionally, installing the shims is typically a time-consuming and expensive process. In some cases, the composite structure may need to be assembled and disassembled several times to measure the shim gap and drill and properly fill the holes. Care must also be taken to ensure that shims fabricated to specific dimensions and shapes to fill a specific gap are used to fill the correct gap and not accidentally used to fill an incorrect gap.

[0005] Some manufacturing processes introduce additional delays at other stages. For example, efforts have been made to reduce the use of shims by custom-manufacturing the spars and spar caps after the skin panels are fabricated, once the as-manufactured dimensions of the skin panels are known. This approach can result in a buildup of inventory of wing skin panels during the manufacture of the spars and spar caps. Wing skin panels are large structures that must be properly held and supported during the manufacture of the spars and spar caps to avoid damage to the skin panels, occupying valuable warehouse space for extended periods of time. Even with these efforts, the use of shims may still be required at some sites. Summary of the Invention [Problem to be solved by the invention]

[0006] The methods disclosed herein are intended to reduce and / or eliminate the need for shims when manufacturing assemblies, such as in the manufacture of aircraft wings. [Means for solving the problem]

[0007] A typical method for manufacturing an assembly of at least a first part and a second part generally includes predicting a set of expected manufacturing dimensions for the first part within predetermined tolerances, manufacturing the first part, scanning the first part to determine a set of actual manufacturing dimensions for the first part, and at least commencing manufacturing the second part before the step of scanning the first part is completed. The predicted manufacturing dimensions may be predicted using a historical data set obtained from multiple previously constructed parts and including dimensions of the previous parts, along with data from multiple factors. The second part may be manufactured based on the set of predicted manufacturing dimensions for the first part such that the second part is configured to mate with the predicted first part. However, such a method may save valuable manufacturing time by allowing the second part to be at least partially or fully completed before the first part has been fully scanned (and therefore the actual as-manufactured dimensions of each first part are known). After scanning of the first part is completed (and the as-manufactured dimensions of the first part determined), the set of predicted manufacturing dimensions may be compared to the set of actual manufacturing dimensions to check for any non-conforming deviations between the set of predicted manufacturing dimensions and the set of actual manufacturing dimensions. Such an approach may enable automated manufacturing of the second part (e.g., not requiring fitting to the second part during the first manufacturing stage), and thus may streamline the assembly process overall, as the second part is modified (e.g., fitted) in areas of any detected non-conforming deviations between the as-manufactured first and second parts.

[0008] Also disclosed is a part scan model that can be configured to create a model of the boundary zone between the first and second mating structures. Such a model may include an original large-scale scan-based layer of the entire boundary zone as initially fabricated and one or more small-scale scan layers interleaved with the large-scale scan-based layer, the small-scale scan layers corresponding to localized portions of the boundary zone that have been modified since initial fabrication. [Brief explanation of the drawings]

[0009] [Figure 1] 1 is a perspective view of a prior art example of an aircraft wing structure; [Figure 2] FIG. 1 is a schematic flow chart diagram illustrating an exemplary method for manufacturing an assembly according to the present disclosure. [Figure 3] 1 is an example of a portion of an assembly that can be manufactured according to the methods of the present disclosure, shown partially exploded; DETAILED DESCRIPTION OF THE INVENTION

[0010] FIG. 2 generally illustrates a flowchart representing a non-limiting, illustrative example of a method 100 according to the present disclosure. In FIG. 2, some steps are indicated by dashed boxes, indicating that such steps may be optional or may correspond to optional examples of a method according to the present disclosure. However, not all methods according to the present disclosure are required to include steps illustrated by solid boxes. The method and steps illustrated in FIG. 2 are not limiting, and as will be understood from the description herein, other methods and steps, including methods having more or fewer steps than those illustrated, are within the scope of the present disclosure.

[0011] The method 100 described herein is broadly applicable to the manufacture of assemblies of two or more parts, such as the combination of a first part and a second part. While the method is, of course, broadly applicable to the manufacture of assemblies of multiple parts, for clarity, it will first be described with reference to the combination of a first part (e.g., a wing skin panel for an aircraft wing) and a second part (e.g., a rib for an aircraft wing). Figure 3 shows a partially disassembled partial aircraft wing assembly 20, which will be referenced in describing the method 100 of Figure 2. The wing assembly 20 includes a skin panel 22 (also sometimes referred to as a wing panel skin, wing skin panel, panel skin, wing skin, or wing panel), a first rib 24, and first and second spar caps 26a, 26b. When completed, the wing assembly 20 will include 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, as will be appreciated by those skilled in the art. 3, the spar caps 26a, 26b are generally secured to the skin panel 22 along a longitudinal boundary band 28 extending longitudinally along the skin panel 22, while the first rib 24 is secured to the skin panel 22 along a transverse boundary band 30 extending transversely across the skin panel 22. Any additional ribs secured to the skin panel 22 may be secured to additional transverse boundary bands spaced apart from the transverse boundary band 30. In examples relating to the assembly of other structures other than the illustrated aircraft wing, the first and second components will widely have boundary bands that differ in number, location, spacing, shape, size, and / or orientation, although these examples do not depart from the scope of this disclosure.

[0012] The disclosed method 100 includes predicting 102 a set of expected manufacturing dimensions for a first part, e.g., a set of expected manufacturing dimensions for a skin panel 22. While a manufactured part will have a set of designed or intended dimensions, it is understood that the actual constructed part may vary slightly from the designed dimensions. For example, while a manufacturing process may intend to produce multiple fasteners, each 1.00 centimeter (cm) long, the actual fasteners manufactured may have different lengths, e.g., between 0.95 and 1.05 cm. Such variations in complexity are also compounded when parts have complex geometries and shapes. Thus, the set of expected manufacturing dimensions may be referred to as an expected range of manufacturing dimensions. In other words, a predetermined and / or acceptable tolerance range (i.e., a range of acceptable specifications, variations, and / or dimensions) may be predicted 102 based on a historical data set of multiple previously constructed parts. For example, in the example of a skin panel 22, the expected manufacturing dimensions may be predicted 102 based on a historical data set of multiple previously constructed skin panels (as opposed to only looking at the design dimensions of the skin panel). Predicting expected manufacturing dimensions at 102 broadly includes characterizing the interface where the parts interlock.

[0013] The step of predicting the expected manufacturing dimensions at 102 is typically based on multiple independent factors and not simply an average of the historical dimensions of constructed parts. For example, the prediction may be based on the source of material used in manufacturing the part, the specific lot of material used, the atmospheric and / or site environmental conditions in which the part is manufactured, the type of bagging or other work material used in layup, the type of prepreg material used, the end effector (or its settings) used in manufacturing, the feed rate of the tape laying head that applies the fiber tape, and / or the temperature and / or pressure used in the autoclave when manufacturing the part. Often, such data related to the manufacturing of parts is recorded, and the methods disclosed herein may utilize any available information to predict the dimensions of future parts. The step of determining the expected manufacturing dimensions at 102 may include statistical process control and / or may utilize analytics and / or machine learning in some methods 100. Such predictive models may take multiple inputs (e.g., some or all of the factors listed above) to predict the dimensions of manufactured parts. For example, a predictive model may predict dimensions for a part manufactured using materials from a first supplier that may differ from dimensions predicted by the predictive model for the same part manufactured using materials from a second supplier. Using the disclosed methods 100, the first part may be manufactured more cheaply. For example, in some methods 100, the step of predicting predicted dimensions at 102 may provide a sufficiently accurate prediction even when a less accurate (and less expensive) manufacturing technique is used for the first part.

[0014] Determining the predicted dimensions at 102 may be performed before, concurrently with, and / or after manufacturing the first part (e.g., skin panel 22) at 104. Manufacturing the first part at 104 may include steps such as trimming one or more layers of the first part, drilling one or more holes through one or more layers of the first part, and / or laying up multiple fiber-reinforced polymer layers. The first part may be scanned at 106 to measure and determine a set of actual manufactured dimensions of the as-built first part. Typically, scanning the first part at 106 involves performing a wide-area scan of the entire part, although in some examples, only a portion of the first part may be scanned at 106. For example, scanning the first part at 106 may be performed by laser scanning, optical scanning, photography, metrology scanning, and / or any other suitable technique. In some examples, scanning the first part at 106 may include placing reference markers on the first part and taking multiple photographs or images of the first part (e.g., using a large camera stage and / or hanging and / or moving the camera relative to the first part) so that the images are stitched together to form a composite scan of the first part using the reference markers as reference points for combining the images. Generally, scanning the first part at 106 includes, at a minimum, scanning any boundary bands where a second part is engaged with the first part. For example, in the case of a skin panel 22, scanning the first part at 106 may include scanning one or more longitudinal boundary bands 28 at 108 (e.g., scanning spar boundaries) and / or scanning one or more transverse boundary bands 30 at 110 (e.g., scanning rib locations). Scanning the first part at 106 may be performed for quality control, recording, learning and / or future analytics purposes.

[0015] Method 100 also includes manufacturing one or more second parts (and / or one or more additional parts) of the assembly at 112, which may at least begin before completing the step of scanning the first part at 106. For example, manufacturing the second part at 112 may be at least partially performed during the scanning of the first part at 106 (e.g., before completing the scanning of the first part) or may be completed entirely. While counterintuitive, this sequencing can improve manufacturing time and efficiency. In the example described herein, manufacturing the second part 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. Based on the predicted manufacturing dimensions of the first part, the second part (or parts) may be manufactured at 112 such that the second part is configured to mate with the first part. In other words, the second part may be manufactured at 112 before the full as-manufactured dimensions of the first part are known, which may allow for automated manufacturing of the second part, as opposed to waiting until the as-manufactured dimensions of the first part are determined before beginning physical manufacturing of the second part.

[0016] Upon completion of scanning the first part at 106, the actual manufacturing dimensions of the first part as manufactured (as determined by scanning the first part at 106) are compared at 118 with the predicted manufacturing dimensions of the first part. Such a comparison may detect any non-conforming deviations between the actual manufactured dimensions and the predicted manufactured dimensions. Criteria for non-conforming deviations may be predetermined based on desired tolerances and tolerances. For example, a non-conforming deviation may be defined as an area where each actual as-manufactured dimension exceeds a maximum dimension in a range of predicted dimensions or is less than a minimum dimension in a range of predicted dimensions. Additionally, the actual manufacturing dimensions of the first part may be added to a historical dataset at 120 for use in predicting dimensions of subsequently manufactured first parts (e.g., for future data analytics analysis of future parts).

[0017] At 122, the first and second parts (along with any additional parts) are joined or combined to form an assembly. For example, the skin panel 22, first rib 24, and spar caps 26a, 26b ( FIG. 3 ) are combined with additional parts to form the aircraft wing assembly 20. In some methods 100, the parts can be mated (e.g., brought together and engaged at a boundary band) without requiring metrology or scanning of the second part (e.g., a ladder structure of ribs and spars) before engaging the second part with the first part (e.g., a skin panel).

[0018] In methods 100 that include a third part (and / or additional second parts and / or additional other parts), the third part may be manufactured and mated to the first part, again based on a set of predicted manufacturing dimensions for the first part. For example, the second part (e.g., first rib 24) and the third part (e.g., first spar cap 26a) may each be manufactured according to the predicted dimensions of the skin panel 22 before the skin panel 22 is fully scanned. In some methods 100, in contrast to the prior art, the manufacturing of each part of the assembly may be decoupled. For example, the spar caps 26a, 26b may be initially manufactured to the same data set each time, without relying on the as-manufactured dimensions of the rib and skin panel when manufacturing the spar caps 26a, 26b.

[0019] In some methods 100, comparing the set of predicted dimensions to the set of actual dimensions at 118 results in the detection of one or more non-conforming deviations at 124. In such methods 100, a repair may be performed at 126 to address or eliminate the non-conforming deviations detected at 124. For example, the first part, the second part, and / or the third part may be modified so that the actual dimensions are within the set of predicted ranges of dimensions, i.e., so that the parts engage and fit together without requiring the placement of shims between the parts (e.g., so that the fit between the first part and the second part is within a predetermined tolerance). In the example of an aircraft wing assembly, a non-conforming deviation may be detected where the as-manufactured dimensions of the parts result in unacceptable interface gaps 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 the installation of shims at the non-conforming interfaces. However, rather than creating shims for such mismatches, the disclosed methods may locally repair the components to eliminate the need for shims. For example, the skin panel 22 may be trimmed, “shaved,” and / or machined to a specified dimension (e.g., removing material), 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 can mate and engage 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 additional components. However, as used herein, the phrase “adding material” does not include inserting shims into the assembly. For example, the removed material may be added to the components in the areas of mismatches and then machined to a specified dimension.

[0020] Restoring the component at 126 may include determining how much material to add to or remove from one or more contact surfaces. For example, restoring the component at 126 may include adding or removing material from the skin panel 22 at one or more different locations along a given transverse boundary band 30 and / or at one or more different locations along a given longitudinal boundary band 28. Additionally or alternatively, restoring the component at 126 may include adding or removing material from a contact surface of the first rib 24 configured to engage with the transverse boundary band 30 and / or adding or removing material from a contact surface of the spar caps 26a and / or 26b configured to engage with the longitudinal boundary band 28. In other words, repairing the parts at 126 may include modifying one or more respective localized regions of the first part, one or more respective localized regions of the second part, and / or one or more respective localized regions of the third part, where the modified localized regions correspond to regions that engage one or more other parts in the detected areas of mismatch. Because the second and third parts are manufactured according to predicted dimensions in method 100 (as opposed to according to designed or intended dimensions), machining the parts for repair at 126 is generally faster and more efficient than creating shims using traditional methods.

[0021] After repairing the part at 126, one or more local areas may be rescanned at 128 in the areas of the non-conforming deviations detected at 124 (e.g., in the local areas where corrections and / or repairs were made at 126). Rescanning at 128 may include rescanning the first part at the boundary locations corresponding to the non-conforming deviations and / or rescanning the assembly of parts in the areas of the non-conforming deviations. Being able to perform the local rescan of the repaired part at 128 without having to place the part on a large gantry advantageously reduces costs in some methods 100. If non-conforming deviations continue to be detected at 124 as a result of the local scan at 128, additional repairs may be performed at 126, and so on, until the finished part is assembled within tolerances. The data (e.g., dimensions) from the local scan at 128 may be digitally inserted (sometimes referred to herein as digitally plugged) into the set of actual manufacturing dimensions at 130 such that the set of actual manufacturing dimensions is accurate to the final as-manufactured part and is then added to the historical data set at 120. For example, reference coordinates of the first part may be used during the rescan at 128 and / or the scan at 106 to facilitate inserting the data from the rescan at 128 into the large scale scan of the first part.

[0022] Thus, method 100 may include creating a part scan model as part of the historical data set, the part scan model including a scan of an interface band between a first mating structure and a second mating structure, such as the transverse and / or longitudinal interface band between a skin panel and a rib and / or spar of the aircraft wing assembly of FIG. 3. In one example, the first mating structure may be a spar interface (e.g., longitudinal interface band 28) of a skin panel for the aircraft wing, and the second mating structure may be a spar cap (e.g., spar caps 26a, 26b) configured to engage with the skin panel. Additionally or alternatively, the first mating structure may be a rib (e.g., transverse interface band 30) of a skin panel for the aircraft wing, and the second mating structure may be a rib (e.g., first rib 24) configured to engage with the skin panel. The part scan model may include an original large-scale scan-based layer of the entire boundary band (e.g., the entire longitudinal boundary band 28 and / or the entire transverse boundary band 30) as initially manufactured, and one or more small-scale scan layers that are interleaved with the large-scale scan-based layer. The small-scale scan layers correspond to localized portions of the one or more boundary bands that have been modified after initial manufacture, such as by repairs or modifications performed at 126 in response to detected misalignments that result in non-conformance.

[0023] Such methods 100 according to the present disclosure can provide numerous advantages in manufacturing, particularly in the manufacture of large assemblies such as aircraft wings. By beginning production of a second part at 112 before the scanning of a first part is complete at 106, factory inventory can be reduced by reducing the time required to store skin panels before assembly with other parts. This reduction in required storage, as well as the reduction in manufacturing time due to not having to form, mold, and install multiple shims into an assembly, may contribute to cost reductions for a given manufacturing process. Use of some methods 100 may reduce the amount of rework required and / or improve the overall part quality of the finished part compared to conventional techniques for installing shims. Again, by attempting to eliminate the need to install shims into an assembly, use of methods 100 may increase manufacturing speed (e.g., throughput) and / or reduce the complexity of manufacturing logistics.

[0024] Non-limiting illustrative examples of the inventive subject matter of the present disclosure are set forth in the enumerated paragraphs below.

[0025] A1. A method of manufacturing an assembly comprising a first part and a second part, the method comprising: predicting a set of expected manufacturing dimensions within predetermined tolerances for a first part using a historical data set of a plurality of previously built parts; manufacturing a first part; scanning the first part to determine a set of actual manufacturing dimensions of the first part; manufacturing a second part, the second part being manufactured based on the set of predicted manufacturing dimensions of the first part such that the second part is configured to mate with the first part, the manufacturing of the second part being performed at least in part before the scanning of the first part is completed; comparing the set of predicted manufacturing dimensions with the set of actual manufacturing dimensions to check for any non-conforming deviations between the set of predicted manufacturing dimensions and the set of actual manufacturing dimensions; A method comprising:

[0026] A1.1. The method of paragraph A1, wherein the step of manufacturing the second part is completed before the step of scanning the first part is completed.

[0027] A1.2. The method of paragraph A1 or A1.1, wherein scanning the first part comprises performing a wide-area scan of the entire first part.

[0028] A1.3. The method of any one of paragraphs A1-A1.2, further comprising adding actual manufacturing dimensions to the historical data set.

[0029] A2. The method of any one of paragraphs A1 through A1.3, further comprising joining a second part to the first part to form an assembly.

[0030] A3. The method of any one of paragraphs A1-A2, wherein the first component comprises a skin panel for an aircraft wing.

[0031] A4. The method of any one of paragraphs A1 to A3, wherein the second component comprises an aircraft wing / a rib for the aircraft wing.

[0032] A5. The method of any one of paragraphs A1-A4, wherein the second component comprises an aircraft wing / spar cap for the aircraft wing.

[0033] A6. The assembly includes a third part, and the method comprises: manufacturing a third part, the third part being manufactured such that the third part is configured to mate with the first part based on the set of predicted manufacturing dimensions of the first part, the manufacturing of the third part being performed at least in part before the scanning of the first part is completed; connecting a third part to the first part to form an assembly; The method of any one of paragraphs A1 to A5, further comprising:

[0034] A6.1. The method of paragraph A6, wherein the manufacturing of the third part is completed before the step of scanning the first part is completed.

[0035] A7. The method of paragraph A6 or A6.1, wherein the first component comprises an aircraft wing / skin panel for the aircraft wing / skin panel, the second component comprises a rib for the aircraft wing, and the third component comprises a spar cap for the aircraft wing.

[0036] A8. Detecting one or more non-conforming deviations between the set of predicted manufacturing dimensions and the set of actual manufacturing dimensions; modifying the second and / or third parts to address one or more mismatched deviations so that the second and / or third parts engage with the first part when assembled within a predetermined tolerance; The method of any one of paragraphs A1 to A7, further comprising:

[0037] A9. The method of paragraph A8, wherein the modifying step comprises one or more steps selected from the group comprising: removing material from the second part; removing material from the third part; adding material to the second part; and adding material to the third part.

[0038] A9.1. The method of paragraph A8 or A9, wherein the modifying step includes determining how much material to remove from and / or add to one or more respective contact surfaces of the second part and / or one or more respective contact surfaces of the third part.

[0039] A10. The method of any one of paragraphs A8 to A9.1, wherein the modifying step comprises modifying one or more respective local areas of the second part and / or one or more respective local areas of the third part, the local areas corresponding to areas of the detected misalignment that engage with the first part.

[0040] A11. The method of any one of paragraphs A8 to A10, further comprising locally scanning one or more local regions corresponding to the detected mismatch.

[0041] A12. The method of paragraph A11, wherein the step of locally scanning one or more local regions is performed after the step of modifying the second part and / or the third part.

[0042] A13. The method of paragraph A11 or A12, further comprising digitally inserting data from the step of locally scanning one or more local regions into a set of actual manufacturing dimensions.

[0043] A14. The method of any one of paragraphs A1 to A13, wherein the predicting step is influenced by multiple independent factors.

[0044] A14.1. The method of paragraph A14, wherein the independent factors include the source of the material, the particular lot of material, the atmospheric and / or site environmental conditions, the type of bagging or other workpiece material used during layup, the type of prepreg material used, the end effector (or its setting) used during production, and / or the temperature and / or pressure used in the autoclave when producing the first part.

[0045] A15. The method of any one of paragraphs A1 through A14.1, wherein the predicting step comprises statistical process control.

[0046] A16. The method of any one of paragraphs A1 to A15, wherein the step of manufacturing the second part is automated by not machining the second part to fit to a specific feature of the first part.

[0047] A17. The step of manufacturing a first part includes: laying up a plurality of fiber reinforced polymer plies; trimming the layer; drilling a plurality of holes in the layer; The method of any one of paragraphs A1 to A16, comprising:

[0048] A18. The method of any one of paragraphs A1 to A17, wherein the first component comprises an aircraft wing / a skin panel for the aircraft wing / the skin panel, the skin panel comprising one or more spar interfaces that engage one or more spars, and the skin panel comprising one or more rib locations that engage one or more ribs, and wherein the scanning step comprises scanning the spar interfaces and scanning the rib locations.

[0049] B1. A scan of a boundary zone between a first mating structure and a second mating structure, the scan comprising: A prototype large-scale scan base layer of the entire boundary zone during the first production, one or more small-scale scan layers inserted into the large-scale scan base layer, the small-scale scan layers corresponding to localized portions of the boundary zone that have been modified after initial fabrication; equipped with, scan A part scan model comprising:

[0050] B2. The part scanned model of paragraph B1, wherein the first mating structure comprises a spar interface of a skin panel for an aircraft wing, and the second mating structure comprises a spar that engages with the skin panel.

[0051] B3. The part scanned model of paragraph B1, wherein the first mating structure comprises a rib portion of a skin panel for an aircraft wing, and the second mating structure comprises a rib that engages with the skin panel.

[0052] D1. Use of the method of any one of paragraphs A1 to A18 to join multiple parts together.

[0053] D2. Use of the method of any one of paragraphs A1 through A18 to manufacture an aircraft wing.

[0054] D3. Use of the method of any one of paragraphs A1 to A18 to connect one or more ribs to a skin panel.

[0055] D4. Use of the method of any one of paragraphs A1 to A18 to join one or more spars and / or spar caps to a skin panel.

[0056] As used herein, the terms "selectively" and "selectively" when referring to modifying the operation, movement, configuration, or other behavior of one or more components or features of a device means that the particular operation, movement, configuration, or other behavior is the direct or indirect result of user manipulation of an aspect or one or more components of the device.

[0057] As used herein, the terms "adapted" and "configured" mean that an element, component, or other protected subject matter is designed and / or intended to perform a certain function. Thus, use of the terms "adapted" and "configured" should not be interpreted to mean that a given element, component, or other protected subject matter is only "capable" of performing a certain function, but rather that the element, component, and / or other protected subject matter is specifically selected, created, implemented, utilized, programmed, and / or designed for the purpose of performing the function. It is within the scope of this disclosure for an element, component, and / or other protected subject matter described as adapted to perform a particular function to be described as configured to perform this function in addition to, or instead of, this description, and vice versa. Similarly, protected subject matter described as configured to perform a particular function may be described as operative to perform this function in addition to, or instead of, this description.

[0058] The phrase "at least one," as used herein in connection with a list of one or more things, should be understood to mean at least one thing selected from any one or more of the things in the list, but does not necessarily include at least one of every thing specifically listed in the list of things, and does not exclude combinations of things in the list. This definition allows for the appropriate presence of things other than those specifically identified in the list of things to which the phrase "at least one" relates, whether or not related to those specifically identified. 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") refers, in one embodiment, to at least one A (including, as appropriate, more than one A) in the absence of B (and, as appropriate, other than B); in another embodiment, to at least one B (including, as appropriate, more than one B) in the absence of A (and, as appropriate, other than A); and, in yet another embodiment, to at least one A (including, as appropriate, more than one A) and at least one B (including, as appropriate, more than one B) (and, as appropriate, other than A). In other words, the terms "at least one," "one or more," and "and / or" are open-ended expressions that are both conjunctive and disjunctive in operation. For example, each of the expressions "at least one of A, B, and C," "at least one of A, B, or C," "one or more of A, B, and C," "one or more of A, B, or 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 any of the above in combination with at least one other, as appropriate.

[0059] The various disclosed elements of the apparatuses and the various disclosed steps of the methods disclosed herein are not required for all apparatuses and methods according to the present disclosure; the present disclosure includes all novel and non-obvious combinations and subcombinations of the various elements and steps disclosed herein. Moreover, one or more of the various elements and steps disclosed herein may define independent subject matter that is distinct and separate from the disclosed apparatus or method as a whole. Thus, such inventive subject matter need not relate to the particular apparatuses and methods explicitly disclosed herein; such inventive subject matter may be useful in apparatuses and / or methods not explicitly disclosed herein.

[0060] As used herein, the words "for example," "for example," and / or the simple phrase "example," when used in connection with one or more components, features, details, structures, embodiments, and / or methods according to the present disclosure, are intended to indicate that the described components, features, details, structures, embodiments, and / or methods are non-limiting, illustrative examples of the components, features, details, structures, embodiments, and / or methods according to the present disclosure. As such, the described components, features, details, structures, embodiments, and / or methods are not intended to be limiting, essential, or exclusive / exhaustive, and other components, features, details, structures, embodiments, and / or methods, including structurally and / or functionally similar and / or equivalent components, features, details, structures, embodiments, and / or methods, are also within the scope of the present disclosure. [Explanation of symbols]

[0061] 10 aircraft wing 12 Ribs 14 digits 16 Skin Panels 20-part aircraft wing assembly 22 Skin Panel 24 First Rib 26a First digit cap 26b Second digit cap 28 Longitudinal boundary zone 30 Short-side boundary strip

Claims

1. 1. A method for generating a part scan model, comprising: scanning a boundary zone of the first mating structure that is engaged by the second mating structure to generate a prototype large scale scan-based layer of the entire boundary zone; comparing a set of predicted manufacturing dimensions of the first mating structure with a set of actual manufacturing dimensions of the first mating structure determined by scanning a boundary zone, and checking for mismatches between the set of predicted manufacturing dimensions and the set of actual manufacturing dimensions; detecting one or more non-conforming deviations between the set of predicted manufacturing dimensions and the set of actual manufacturing dimensions; modifying localized portions of the boundary band to address detected non-conforming deviations; scanning localized portions of the modified boundary zone to generate one or more small-scale scan layers that are interpolated into the original large-scale scan base layer; A method comprising:

2. 2. The method of claim 1, wherein the first mating structure comprises a spar interface of a skin panel (22) for an aircraft wing, and the second mating structure comprises a spar that engages the skin panel (22).

3. 3. The method of claim 1, wherein the first mating structure comprises a rib portion of a skin panel (22) for an aircraft wing, and the second mating structure comprises a rib that engages with the skin panel (22).

4. A method according to any one of claims 1 to 3, further comprising the steps of: correcting the local portion of the boundary band comprises repairing the local portion of the boundary band; and repairing the local portion of the boundary band comprises repairing the one or more misalignments.

5. The method of any one of claims 1 to 4, further comprising interleaving the small-scale scan layer with the original large-scale scan base layer to create a part scan model.

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