Wing components for aircraft, their manufacturing methods and loading methods

Abstract

A wing assembly includes a central wing structure and a pair of outer wing structures. The central wing structure includes a central wing forward sparb, a central wing rear sparb, and engine mounting locations on each side of the wing centerline. Each outer wing structure includes an outer wing forward sparb and an outer wing rear sparb, configured to connect to the central wing forward sparb and central wing rear sparb, respectively, to define wing joints connecting the outer wing structures to the central wing structure. The central wing structure is configured such that the sparb ends of the central wing forward sparb and central wing rear sparb are located at each wing joint no more inward than the engine centerline associated with the engine mounting location, and no more outward than 10 percent of the distance between the engine centerline and the wing centerline.

Description

Technical Field

[0001] This disclosure generally relates to structural devices for aircraft wings, and more specifically, to composite wing assemblies having wing joints for connecting outer wing structures to central wing structures. Background Technology

[0002] For many aircraft, such as commercial transport aircraft made of composite materials, wing assemblies typically consist of a center wing box and a pair of wings. The center wing box is attached to the fuselage. Each of the opposite sides of the center wing box has a fuselage side (SOB) connector located near the fuselage side to attach each wing to the center wing box.

[0003] SOB joints are structurally complex and time-consuming to manufacture. This complexity stems in part from the significant differences in design standards and load conditions between the inner and outer sections of each wing. For example, the inner section of each wing must be relatively rigid and must include multiple locations where the skin panels are locally thickened to accommodate high-load joints, such as SOB joints, landing gear attachments, and engine attachments. These thickened skin panel locations present manufacturing challenges and increase the structural mass of the aircraft. Conversely, the outer sections of each wing typically do not have high-load joints and are preferably flexible to meet aeroelastic performance requirements.

[0004] Stringers are typically attached to the wing skin panels to increase their bending stiffness. In conventional wing assemblies, the stringers in each wing extend continuously from the SOB joint to the wingtip. While stringers are effective in providing relatively high stiffness on the inner portion of each wing, their structural efficiency decreases on the outer portion and can lead to high dynamic loads due to their relatively large mass and relatively high stiffness.

[0005] It can be seen that there is a need in the field for a composite wing assembly that avoids the aforementioned challenges associated with the manufacture, assembly, maintenance, and increased structural mass of conventional composite wing assemblies. Summary of the Invention

[0006] The aforementioned requirements related to wing assemblies are specifically addressed by this disclosure, which provides a wing assembly for an aircraft. The wing assembly includes a central wing structure and a pair of outer wing structures. The central wing structure includes a pair of outer wing ends and a pair of engine mounting positions located on opposite sides of the wing centerline. Additionally, the central wing structure includes a central wing leading sparb and a central wing trailing sparb, each having a sparb end at each of the outer wing ends. Each of the outer wing structures includes an outer wing leading sparb and an outer wing trailing sparb, configured to connect to the central wing leading sparb and the central wing trailing sparb, respectively, to define a wing joint connecting the outer wing structures to the central wing structure. The center wing structure is designed such that the spar ends of the front and rear spars of the center wing are located at each wing joint no more inward than the engine centerline relative to the engine mounting location, and no more than 10 percent of the distance between the engine centerline and the wing centerline, further outward from the engine centerline.

[0007] Furthermore, a composite wing assembly having the central wing structure described above is disclosed. The wing assembly includes a pair of outer wing structures, each outer wing structure having a front wing sparb and a rear wing sparb. The front and rear wing spars are configured to connect to the front and rear wing spars of the central wing, respectively, to define wing joints connecting the outer wing structures to the central wing structure. The outer wing structures also include an upper outer wing skin panel and a lower outer wing skin panel, each comprising an area layer having a panel thickness constant along the wingspan direction. The upper and lower outer wing skin panels are connected to the front and rear wing spars. Additionally, the outer wing structures include one or more slats bonded to the inner surface of at least one of the upper and lower skin panels and extending along the wingspan direction. Each slat consists of a stack of slat layers narrower than the area layer. As described above, the center wing structure is configured such that the spar ends of the front and rear spars of the center wing are located at each wing joint no more inward than the engine centerline associated with the engine mounting location, and no more outward than 10 percent of the distance between the engine centerline and the wing centerline.

[0008] Additionally, a composite wing assembly having the central wing structure described above is disclosed. The wing assembly includes a pair of outer wing structures constructed as described above. In addition to replacing one or more slats bonded to the upper and / or lower skin panels, each outer wing structure includes one or more outer wing stringers connected to the upper and / or lower outer wing skin panels and extending in the spanwise direction. Each outer wing stringer has a cap-shaped cross-section with a stringer cap containing one or more 0-degree layer strips interwoven within the stringer cap. A composite wing assembly having the central wing structure described above is also disclosed. The wing assembly includes a pair of outer wing structures constructed as described above. In addition to replacing slats or outer wing stringers bonded to the upper and / or lower outer wing skin panels, each outer wing structure has an upper outer wing panel assembly and an outer wing lower panel assembly, each extending in the spanwise direction and connected to the outer wing's front and rear wing spars. Each panel assembly includes an outer skin panel consisting of area layers and having a constant panel thickness along the wingspan. Additionally, each panel includes a cap-shaped section panel attached to the inside of the outer skin panel and having multiple generally parallel cap-shaped sections, each extending in the wingspan direction.

[0009] Additionally, a method for manufacturing a wing assembly is disclosed. This method includes attaching the inner ends of the outer wings of each of a pair of outer wing structures to the opposite outer ends of the center wing of a central wing structure, thereby defining a pair of wing joints connecting the outer wing structures to the central wing structure. As described above, the central wing structure has a front wing spade and a rear wing spade, each having a spade end at each of the outer ends of the central wing. The wing assembly includes engine mounting positions on each of the opposite sides of the wing centerline, and each of the opposite spade ends of the front and rear wing spades is located at each wing joint no more inward than the engine centerline associated with the engine mounting position, and no more outward than 10 percent of the distance between the engine centerline and the wing centerline.

[0010] A method for manufacturing a wing assembly is also disclosed, comprising bonding one or more slats to the inner surface of at least one of an upper wing skin panel and a lower wing skin panel in each of a pair of outer wing structures. Each slat extends in the spanwise direction and consists of a stack of slat layers. The upper and lower wing skin panels each comprise an area layer having a constant panel thickness along the spanwise direction. The slat layer has a width narrower than the area layer. The method includes attaching the upper and lower wing skin panels to the outer wing forward and aft wing spars of the outer wing structure. Additionally, the method includes attaching the outer wing forward and aft wing spars of each of the outer wing structures to the center wing forward and aft wing spars of the center wing structure, respectively, to define a pair of wing joints at opposite outer ends of the center wing. As described above, the center wing structure has engine mounting locations on each opposite side of the wing centerline. The spar ends of the center wing's fore spar and aft spar are located at each wing joint no more inside the engine centerline relative to the engine mounting location, and no more outside the engine centerline than 10 percent of the distance between the engine centerline and the wing centerline.

[0011] Additionally, a manufacturing method and wing assembly are disclosed. Similar to the method described above, except that instead of bonding slats to the outer wing upper skin panel and the outer wing lower skin panel, this method includes bonding one or more outer wing stringers to the outer wing upper skin panel and / or to the outer wing lower skin panel. Each outer wing stringer extends in the spanwise direction and includes multiple stringer layers. Each outer wing stringer has a stringer cap comprising one or more 0-degree layer strips. The outer wing upper skin panel and the outer wing lower skin panel each include an area layer having a panel thickness that is constant along the spanwise direction.

[0012] Furthermore, a method for manufacturing a wing assembly is disclosed, comprising bonding a hat-shaped section panel to an outer skin panel of at least one of an upper outer wing panel assembly and an lower outer wing panel assembly. The hat-shaped section panel has a plurality of generally parallel hat-shaped sections, each extending in the wingspan direction. The outer skin panel includes an area layer having a panel thickness with a constant thickness along the wingspan direction. The method further includes attaching the upper outer wing panel assembly and the lower outer wing panel assembly to the outer wing forward spars and outer wing aft spars of an outer wing structure. Additionally, the method includes attaching the inner ends of the outer wings of each of a pair of outer wing structures to the opposite outer ends of the center wing of a center wing structure, respectively, to define a pair of wing joints. As described above, the spars ends of the center wing forward spars and center wing aft spars are located at each wing joint no more inward than the engine centerline associated with the engine mounting location, and no more outward than 10 percent of the distance between the engine centerline and the wing centerline.

[0013] Additionally, a method for loading a wing assembly is disclosed. The method includes providing a wing assembly having an outer wing structure, each outer wing structure connected to a central wing structure, as described above. The method also includes placing the wing assembly under static ground load and under dynamic ground load.

[0014] The features, functions and advantages discussed can be implemented independently in various embodiments of this disclosure, or can be combined in other embodiments, further details of which can be seen in the following description and figures. Attached Figure Description

[0015] These and other features of this disclosure will become more apparent when referenced to the accompanying drawings, wherein the same numerals always denote the same parts, wherein:

[0016] Figure 1 It is a top view of an example of an airplane;

[0017] Figure 2 This is a top view of an example of a wing assembly having a central wing structure and an outer wing structure, each outer wing structure being attached to the central wing structure at a wing joint outside the engine mounting position located on each side of the central wing structure.

[0018] Figure 3 yes Figure 2 An exploded view of the wing assembly, showing the outer wing structure separate from the central wing structure;

[0019] Figure 4 This is a top view of an example of an outer wing structure that is joined to the center wing structure at the wing joint, and also shows the location of the fuel tank in the wing assembly;

[0020] Figure 5 This is an exploded cutaway perspective view of an example of an outer wing structure having an outer wing transverse rib configured to be mechanically fastened to a center wing transverse rib, and examples of a front wing sparsity splice plate and a rear wing sparsity splice plate are also shown. The front wing sparsity splice plate is configured to connect the center wing front wing sparsity to the outer wing front wing sparsity, and the rear wing sparsity splice plate is configured to connect the center wing rear wing sparsity to the outer wing rear wing sparsity.

[0021] Figure 6 It is along Figure 4 The cross-sectional view of the wing assembly taken by line 6-6 shows the central wing stringers that pass through the central wing ribs of the central wing structure;

[0022] Figure 7 yes Figure 6An enlarged view of a portion of the wing assembly identified by reference numeral 7, showing the central wing stringer passing through the rib cutouts formed in the central wing rib, and also showing the central wing rib, the upper skin panel of the central wing, and the lower skin panel of the central wing connected to the front wing spars of the central wing.

[0023] Figure 8 yes Figure 6 An enlarged view of a portion of the wing assembly identified by reference numeral 8 in the attached figure shows the central wing rib, the upper skin panel of the central wing, and the lower skin panel of the central wing, which are connected to the rear wing spars of the central wing.

[0024] Figure 9 yes Figure 7 A view of an example of a hybrid construction of a central wing front spars having a pair of L-shaped rib chords connected to each other by rib webs;

[0025] Figure 10 yes Figure 8 A view of an example of a center wing rear sparsity, which is constructed similarly to... Figure 9 The center wing forward spars;

[0026] Figure 11 It is along Figure 7 The cross-sectional view taken by line 11-11 in the figure shows an example of a center wing rib having an upper rib flange and a lower rib flange respectively connected to the upper skin panel and the lower skin panel of the center wing.

[0027] Figure 12 It is along Figure 7 The cross-sectional view taken by line 12-12 shows an example of a central wing stringer passing through the rib cutouts in the central wing ribs of the upper and lower central wing skin panels.

[0028] Figure 13 yes Figure 11 A view of an example of a central wing rib, wherein the rib web and the rib flange are composed of fabric layers, and the rib flange contains one or more 0-degree layer strips interwoven within the fabric layers;

[0029] Figure 14 It is along Figure 4 The cross-sectional view of the wing assembly taken by line 14-14 shows an example of a hole pattern for receiving multiple mechanical fasteners used to connect the center wing cross rib to the outer wing cross rib.

[0030] Figure 15 yes Figure 14 An enlarged view of a portion of the wing assembly identified by reference numeral 15, showing the central wing transverse rib, the upper skin panel of the central wing, and the lower skin panel of the central wing connected to the central wing front spars.

[0031] Figure 16 This is an enlarged view of a portion of the wing assembly identified by reference numeral 16, showing the center wing transverse rib, the upper skin panel of the center wing, and the lower skin panel of the center wing, which are connected to the rear spars of the center wing.

[0032] Figure 17 It is along Figure 15 The wing assembly section cut by line 17-17 shows an example of a wing joint, showing the central wing transverse rib and the outer wing transverse rib fastened together back to back.

[0033] Figure 18 yes Figure 17 Exploded view of the wing joint;

[0034] Figure 19 yes Figure 18 A cross-sectional view of an example of an outer wing transverse rib;

[0035] Figure 20 yes Figure 18 An enlarged view of a portion of the center wing structure, identified by reference numeral 20, showing an example of a raised area in the lower skin panel of the center wing, located where the center wing stringers terminate at the center wing transverse ribs;

[0036] Figure 21 yes Figure 20 An enlarged view of a portion of the wing assembly identified by reference numeral 21 in the accompanying drawings shows the distribution of shear and pull-out stresses along the bonding line that connects the center wing stringer to the center wing lower skin panel.

[0037] Figure 22 It is along Figure 17 The top view of the wing assembly taken by line 22-22 shows the wing joints and ends of the center wing stringer and the outer wing stringer, which are located on opposite sides of the wing joint, respectively.

[0038] Figure 23 It is along Figure 16 The cross-sectional view taken by line 23-23 shows an example of a rear wing spars splice plate that connects the center wing rear spar to the outer wing rear spar at the wing joint;

[0039] Figure 24 yes Figure 23 A partial exploded view of the wing joint;

[0040] Figure 25 It is along Figure 23 The cross-sectional view of the wing assembly taken by line 25-25 shows an example of a rear wing sparsity splice used to connect the center wing rear wing sparsity to the outer wing rear wing sparsity.

[0041] Figure 26 yes Figure 25 Exploded view;

[0042] Figure 27 yes Figure 24 An enlarged view of a portion of the wing assembly identified by reference numeral 27 shows an example of a raised area in the lower skin panel of the center wing, located where the rear spars of the center wing terminates at the center wing transverse rib.

[0043] Figure 28 This is a partial sectional perspective view of an example of an outer wing structure, showing the outer wing stringers used to reinforce the lower skin panel of the outer wing in the inner portion of the outer wing structure, and the 0-degree strip group used to reinforce the lower skin panel of the outer wing in the outer portion of the outer wing structure.

[0044] Figure 29 This is a top view of an example of an outer wing structure;

[0045] Figure 30 It is along Figure 29 The cross-sectional view taken by line 30-30 shows an example of an outer wing stringer passing through a rib cutout formed in the outer wing rib;

[0046] Figure 31 yes Figure 30 An enlarged view of a portion of the wing assembly identified by reference numeral 31, showing the outer wing transverse ribs connecting the outer wing upper skin panel, the outer wing lower skin panel, and the outer wing front spars.

[0047] Figure 32 yes Figure 30 An enlarged view of a portion of the wing assembly identified by reference numeral 32, showing the outer wing transverse ribs connecting the outer wing upper skin panel, the outer wing lower skin panel, and the outer wing rear spars.

[0048] Figure 33 It is along Figure 29 The cross-sectional view taken by line 33-33 shows the termination of the outer wing stringer at the outer wing rib located near the outer end of the fuel tank of the outer wing;

[0049] Figure 34 yes Figure 33 An enlarged view of a portion of the wing assembly identified by reference numeral 34 in the accompanying drawings shows a partial raised area in the outer wing upper skin panel and the outer wing lower skin panel at the end of the outer wing stringer and at the outer wing front spars.

[0050] Figure 35 yes Figure 33An enlarged view of a portion of the wing assembly identified by reference numeral 35, showing a partial raised area in the outer wing upper skin panel and the outer wing lower skin panel at the end of the outer wing stringer and at the outer wing rear spars.

[0051] Figure 36 It is along Figure 34 The cross-sectional view of the outer wing structure taken by line 36-36 shows the termination of the outer wing stringers at the outer wing ribs located near the outer end of the fuel tank of the wing assembly.

[0052] Figure 37 yes Figure 36 An enlarged view of a portion of the wing assembly identified by reference numeral 37 shows an example of a partial raised area in the lower skin panel of the outer wing at the end of the outer wing stringer.

[0053] Figure 38 It is along Figure 34 The cross-sectional view taken by line 38-38 shows the outer wing forespar extending through the outer wing rib located near the outer end of the fuel tank of the wing assembly.

[0054] Figure 39 yes Figure 38 An enlarged view of a portion of the outer wing structure identified by reference numeral 39 shows a partial raised area in the lower skin panel of the outer wing, located near the outer end of the outer wing rib.

[0055] Figure 40 It is along Figure 29 The cross-sectional view taken by line 40-40 shows the upper skin panel and the lower skin panel of the outer wing, each having one or more 0-degree strip groups, each strip group extending in the wingspan direction to reinforce the skin panel.

[0056] Figure 41 yes Figure 40 An enlarged view of a portion of the outer wing lower skin panel, identified by reference numeral 41, shows an example of a 0-degree strip group comprising 0-degree layer strips interwoven within the area layers that make up the outer wing lower skin panel.

[0057] Figure 42 This is another example of 0-degree layer stripes interwoven within the area layer of the outer wing upper skin panel or the outer wing lower skin panel;

[0058] Figure 43 It is along Figure 29 The wing assembly cross-sectional view taken by line 43-43 shows an example of ply reduction in the 0-degree layer strip of the lower skin panel of the outer wing;

[0059] Figure 44 It is along Figure 29The cross-sectional view taken by line 44-44 shows an example of a 0-degree layer strip interwoven within the lower skin panel of the outer wing at a chord position directly below the outer wing stringer.

[0060] Figure 45 yes Figure 44 An enlarged view of a portion of the outer wing lower skin panel, identified by reference numeral 45, showing the outer wing stringers located directly above the 0-degree layer strips interwoven within the area layer of the outer wing lower skin panel.

[0061] Figure 46 This is a flowchart of a method for manufacturing a wing assembly by attaching an outer wing structure to each of the opposite outer ends of a central wing structure.

[0062] Figure 47 This is a flowchart of a method for loading a wing assembly, which includes a pair of outer wing structures connected to a central wing structure;

[0063] Figure 48 This is a partial sectional perspective view of another example of an outer wing structure with separate slats for reinforcing the upper and lower skin panels of the outer wing;

[0064] Figure 49 yes Figure 48 A top view of an example of an outer wing structure;

[0065] Figure 50 It is along Figure 49 The sectional view along line 50-50 shows examples of multiple strips bonded to the inner surfaces of the upper and lower skin panels.

[0066] Figure 51 yes Figure 50 An enlarged view of a portion of the outer wing structure identified by reference numeral 51 shows an example of the slat-skin section cross-sectional area on each of the upper and lower skin panels of the outer wing, corresponding to the combined slat-skin section bending stiffness of the skin panel portion and the associated slats.

[0067] Figure 52 It is along Figure 49 The cross-sectional view taken by line 52-52 shows an example of a slat passing through a rib cutout formed in the outer wing rib;

[0068] Figure 53 yes Figure 52 An enlarged view of a portion of the outer wing structure identified by reference numeral 53, showing rib cutouts formed in the outer wing ribs;

[0069] Figure 54 yes Figure 51An enlarged view of a portion of the outer wing structure, identified by reference numeral 54, shows an example of a stack of slats whose width is narrower than the area layers of the upper and lower skin panels of the outer wing;

[0070] Figure 55 This is a cross-sectional view of an example of a trapezoidal cross-section strip bonded to the lower skin panel of the outer wing;

[0071] Figure 56 It is along Figure 49 The cross-sectional view of the outer wing structure taken by line 56-56 shows an example of a decreasing layup in a slat layer of one of the slats bonded to the lower skin panel of the outer wing.

[0072] Figure 57 This is a flowchart of a method for manufacturing wing components by connecting a slat-reinforced outer wing structure to a central wing structure.

[0073] Figure 58 It is a flowchart of a method for loading a wing assembly, which includes a pair of slat-reinforced outer wing structures connected to a central wing structure;

[0074] Figure 59 This is a partial sectional perspective view of another example of an outer wing structure, in which the upper and lower skin panels of the outer wing are reinforced by outer wing stringers, each having a hat-shaped cross section;

[0075] Figure 60 yes Figure 59 A top view of an example of an outer wing structure;

[0076] Figure 61 It is along Figure 60 The cross-sectional view taken by line 61-61 shows an example of the upper skin panel and the lower skin panel of the outer wing, each of which is reinforced by the outer wing stringer with a hat-shaped cross section.

[0077] Figure 62 yes Figure 61 An enlarged view of a portion of the outer wing structure identified by reference numeral 62 shows an example of the cross-sectional area of ​​the stringer-skin portion on each of the upper and lower skin panels of the outer wing, which corresponds to the combined stringer-skin portion bending stiffness of the skin panel portion and the associated outer wing stringer.

[0078] Figure 63 It is along Figure 60 The cross-sectional view taken by line 63-63 shows an example of an outer wing stringer passing through a rib cutout formed in the outer wing rib;

[0079] Figure 64 yes Figure 63 An enlarged view of a portion of the outer wing structure identified by reference numeral 64, showing rib cutouts formed in the outer wing ribs for the passage of the outer wing stringers;

[0080] Figure 65 yes Figure 62 An enlarged view of a portion of the outer wing structure identified by reference numeral 65 in the accompanying drawings shows an example of the laminated layer of the outer wing stringers, and also shows 0-degree layer strips interwoven within the stringer cover;

[0081] Figure 66 yes Figure 65 An enlarged view of a portion of the outer wing stringer, identified by reference numeral 66, showing the 0-degree layer strip in the stringer cover;

[0082] Figure 67 yes Figure 65 An enlarged view of a portion of the skin panel and outer wing stringer, identified by reference numeral 67, showing 0-degree layer strips interwoven within the composite layer of the lower skin panel and stringer base laminate;

[0083] Figure 68 It is a cross-sectional view of an example of an outer wing stringer with a stringer base laminate, the stringer base laminate including a 0-degree layer and being connected to a lower skin panel without a 0-degree layer strip.

[0084] Figure 69 This is a cross-sectional view of an example of an outer wing stringer, in which the stringer base laminate and lower skin panel do not have 0-degree layer strips;

[0085] Figure 70 It is along Figure 60 The cross-sectional view of the outer wing structure taken from line 70-70 shows the ply reduction in the 0-degree layer strips of the stringer cover, stringer base laminate, and lower skin panel of the outer wing.

[0086] Figure 71 This is a flowchart of a method for manufacturing wing components by connecting a stringer-reinforced outer wing structure to a central wing structure.

[0087] Figure 72 This is a flowchart of a method for loading a wing assembly, including an outer wing structure reinforced with a pair of stringers, onto a central wing structure;

[0088] Figure 73 This is a partial sectional perspective view of an example of an outer wing structure having an upper outer wing panel assembly and a lower outer wing panel assembly;

[0089] Figure 74 yes Figure 73 A top view of an example of an outer wing structure;

[0090] Figure 75 It is along Figure 74 A cross-sectional view taken from line 75-75;

[0091] Figure 76 yes Figure 75 An enlarged view of a portion of the outer wing structure identified by reference numeral 76, showing the outer skin panel, cap-shaped section panel, foam member, and slats that make up each of the upper and lower outer wing panel assemblies;

[0092] Figure 77 This is an exploded view of the outer wing structure, showing the outer wing ribs, cap-shaped cross-section panels, foam components, slats, and outer wing panels of the upper and lower outer wing panel assemblies.

[0093] Figure 78 It is along Figure 74 The cross-sectional view taken by line 78-78 shows the outer wing rib with a rib cutout, which is used to accommodate the cap-shaped section of the outer wing upper panel assembly and the outer wing lower panel assembly.

[0094] Figure 79 yes Figure 78 An enlarged view of a portion of the outer wing structure identified by reference numeral 79, showing the rib cutouts of the cap-shaped section for accommodating the cap-shaped section panel;

[0095] Figure 80 yes Figure 76 An enlarged view of a portion of the outer wing structure identified by reference numeral 80, showing 0-degree layer strips interwoven in the cap-shaped section cover of the cap-shaped section, and also showing slats on opposite sides of the foam component;

[0096] Figure 81 yes Figure 76 An enlarged view of a portion of the outer wing structure identified by reference numeral 81 shows slats consisting of multiple layers of slats sandwiched between the outer skin panel and the cap-section flange of the cap-section panel;

[0097] Figure 82 yes Figure 80 An enlarged view of a portion of the cap-shaped section panel identified by reference numeral 82, showing the 0-degree layer strips interwoven within the cap-shaped section cover of the cap-shaped section;

[0098] Figure 83 It is along Figure 80 The transverse cross-sectional view of the outer wing upper panel assembly taken from line 83-83 shows the ply reduction in the 0-degree layer strip of the cap-shaped section cover;

[0099] Figure 84 It is along Figure 81The transverse cross-sectional view of the outer wing upper panel assembly taken by line 84-84 shows an example of ply reduction in the lamination layers of the slats;

[0100] Figure 85 This is a flowchart of a method for manufacturing wing components by connecting an outer wing structure reinforced with a hat-shaped cross-section panel to a central wing structure.

[0101] Figure 86 It is a flowchart of a method for loading a wing assembly, including a pair of hat-shaped cross-section panels reinforcing an outer wing structure, onto a central wing structure. Detailed Implementation

[0102] Reference is now made to the accompanying drawings, which illustrate various embodiments of this disclosure. Figure 1 The image shown is a top view of an example of an aircraft 100. The aircraft 100 includes a fuselage 102, wing assemblies 120 coupled to the fuselage 102, and a pair of engines 106 each mounted to the wing assemblies 120. Figures 1 to 5 As shown, the wing assembly 120 includes a central wing structure 150 and a pair of outer wing structures 210. The outer wing structures 210 are configured to be connected to the central wing structure 150. The central wing structure 150 includes a pair of outer wing tips 168 and a pair of engine mounting positions located on opposite sides of the wing centerline 124. Additionally, the central wing structure 150 includes a leading wing sparb 152 and a trailing wing sparb 154, each formed of composite material and having a sparb end 156 at each outer wing tip 168.

[0103] Each of the outer wing structures 210 includes a front wing spade 212 and a rear wing spade 214, each formed of composite material, and configured to connect to the front wing spade 152 and the rear wing spade 154 of the center wing, respectively, to define a wing joint 172 connecting the outer wing structure 210 to the center wing structure 150. The center wing structure 150 is configured such that the spade ends 156 of the front wing spade 152 and the rear wing spade 154 are located at each wing joint 172 no more inward than a vertical plane (not shown) coinciding with the engine centerline 106 associated with the engine mounting position, and no more outward than the maximum outer distance 171 of the wing joint, which is defined as ten percent of the distance between the engine centerline 106 and a vertical plane (not shown) coinciding with the wing centerline 124. For the purposes of this disclosure, when the aircraft 100 is viewed from a top-down perspective, the engine centerline 106 can be approximately parallel to the wing centerline 124, and the distance between the engine centerline 106 and the wing centerline 124 can be measured perpendicular to the respective vertical plane.

[0104] like Figures 2 to 4 As shown, the center wing structure 150 can be configured as a single piece and may include a means for connecting the wing assembly 120 to the fuselage 102. Figure 1 The center wing box 122. In some instances, the center wing forward spars 152 and the center wing aft spars 154 may extend continuously between the outer ends 168 of the center wing. Advantageously, by dividing the wing assembly 120 into three main structural components, including the center wing structure 150 and a pair of outer wing structures 210, each structural component can be configured to adapt to the specific design standards and load conditions it will experience. For example, the center wing structure 150 can be configured as a relatively rigid structure, with a main joint designed to accommodate high loads. This main joint includes the attachment of the wing assembly 120 to the fuselage 102, and the attachment of the landing gear (not shown) and the engine 106 to the wing assembly 120. Conversely, the outer wing structures 210 may not include high-load joints and can therefore be configured as relatively flexible structures to accommodate dynamic loads related to aeroelastic performance requirements. The relatively flexible construction of the outer wing structures 210 can result in a reduction in the structural mass of the wing assembly 120 relative to the structural mass of a conventional wing.

[0105] refer to Figures 2 to 5 The center wing structure 150 may include a plurality of center wing ribs 174 formed of composite materials, each center wing rib extending between the center wing forward spars 152 and the center wing aft spars 154 and positioned at intervals along the spanwise direction of the center wing structure 150. A plurality of center wing nacelles 188 are defined between adjacent pairs of center wing ribs 174. The center wing ribs 174 may include center wing stout ribs 170 located at wing joints 172 at the outer end 168 of each center wing. As described above, the center wing stout ribs 170 at each wing joint 172 may be formed of a metallic material to provide a higher strength capability than the remaining center wing ribs 174, thereby allowing the center wing stout ribs 170 to transfer at least a portion of the load from the outer wing structure 210 to the center wing structure 150, as described in more detail below.

[0106] refer to Figures 4 to 5The center wing structure 150 may further include an upper center wing skin panel 190 and a lower center wing skin panel 192, each formed of a composite material and connected to the center wing forward spars 152, the center wing rear spars 154, and the center wing rib 174. The upper center wing skin panel 190 and the lower center wing skin panel 192 may be composed of an area layer 302, which can be described as a composite layer 300, extending continuously from the center wing forward spars 152 to the center wing rear spars 154 in the chord direction and continuously in the wingspan direction. For example, the upper center wing skin panel 190 and / or the lower center wing skin panel 192 may extend at least from the fuselage 102 (… Figure 1 The position on the side extends to the center wing transverse rib 170.

[0107] refer to Figure 4 The wing assembly 120 may include one or more fuel tanks 128 for supplying fuel to the engine 106. Figure 1 The fuel tanks 128 are supplied with fuel. The internal structures of the center wing structure 150 and / or the outer wing structure 210 (e.g., spars, wing ribs, and skin panels) may define walls for one or more fuel tanks 128. Alternatively, each fuel tank 128 may be a separate structure (not shown) mounted within the wing assembly 120. Each wing assembly 120 may include or define multiple fuel tanks 128. In the example shown, the center wing structure 150 may include a main tank 130 on each side of the wing centerline 124. Figure 4 Additionally, each outer wing structure 210 may include another main housing 130. Figure 4 ) and ventilation box 132 located outside the main box 130 ( Figure 4 In this disclosure, the outermost fuel tank 128 in each outer wing structure 210 has an outer end 134 that may define the end of an outer wing stringer 272, which may be included in the outer wing structure 210 as described below.

[0108] As described above, components of the center wing structure 150 and the outer wing structure 210 are formed of composite materials. Advantageously, structures formed of composite materials exhibit improved corrosion resistance and fatigue strength compared to metallic materials (e.g., aluminum) used in conventional wings. The composite materials used to manufacture components of the center wing structure 150 and the outer wing structure 210 (e.g., spars, wing ribs, skin panels) may consist of composite layers 300 (e.g., prepreg) of a fiber-reinforced polymer matrix material. Each composite layer 300 may have a thickness of 0.005 inches or greater. The polymer matrix material may be a thermosetting resin or a thermoplastic resin. The fibers may be carbon fibers, or the fibers may be formed from alternative materials such as glass, boron, aramid, ceramics, or other non-metallic or metallic materials.

[0109] For certain components, such as the upper and / or lower skin panels of the center wing structure 150 and / or the outer wing structure 210, the composite layer 300 may primarily comprise unidirectional layers 306, each containing a plurality of generally parallel fibers oriented in a single direction. The unidirectional layers 306 may include 0-degree layers, +45-degree layers, -45-degree layers, and 90-degree layers, and / or composite layers 300 having one or more other fiber orientations. Other components, such as the front spar, rear spar, and wing ribs of the center wing structure 150 and / or the outer wing structure 210, may primarily consist of fabric layers 304. In this disclosure, fabric layers 304 may be defined as having multidirectional fibers. For example, fabric layers 304 may include a woven arrangement of bidirectional fibers oriented perpendicular to each other. Advantageously, manufacturing components from fabric layers 304 can reduce or avoid warping (e.g., spring-in) and / or wrinkling, which may occur if the component is formed from unidirectional layers 306.

[0110] In some configurations, the upper center wing skin panel 190 and / or the lower center wing skin panel 192 may include one or more partially raised areas 200 (e.g., Figure 20 As described in more detail below, the padding area 200 may be positioned where the upper center wing skin panel 190 and / or the lower center wing skin panel 192 overlap and are attached (e.g., mechanically fastened and / or bonded) to the locations of the center wing forward spars 152 and the center wing aft spars 154. The panel thickness 194 of the area layer 302 (e.g., unidirectional layer 306) in the skin panel... Figure 20 This eliminates the additional thickness caused by the composite layer 300 that makes up the padding region 200. The panel thickness 194 of the area layer 302 can be substantially constant (e.g., within 20% of each other at any spanwise location), which simplifies the manufacture of the upper center wing skin panel 190 and / or the lower center wing skin panel 192 by eliminating the need to incorporate a ply drop in the area layer 302 as a way to gradually reduce the bending stiffness of the wing assembly 120 along the outward direction.

[0111] Although Figures 1 to 5Each wing joint 172 is shown defining a straight line from the leading edge to the trailing edge of the wing, and each wing joint 172 is shown as substantially perpendicular (e.g., within 20 degrees) to the center wing forward spars 152 and / or the center wing aft spars 154, but the wing assembly 120 may be configured such that each wing joint 172 has a non-linear shape (not shown) from the leading edge to the trailing edge of the wing. For example, the portion of the wing joint 172 from the leading edge to the forward spars may be parallel to the centerline of the aircraft 100. The wing joint 172 can be described as an installation joint, wherein each outer wing structure 210 can be removed and replaced at a maintenance facility, unlike permanent joints manufactured at the aircraft production facility. In this respect, each wing joint 172 may be configured such that the outer wing structure 210 can be removed via the removal of a plurality of mechanical fasteners 298 ( Figure 17 It is removed from the center wing structure 150, as described below.

[0112] refer to Figures 4 to 5 As described in more detail below, each outer wing structure 210 may include a plurality of outer wing ribs 234, each outer wing rib being formed of a composite material (e.g., primarily fabric layer 304) and extending between the outer wing's forward spar 212 and rear spar 214. The outer wing ribs 234 may be positioned at intervals along the wingspan and may define a plurality of outer wing nacelles 236 between adjacent outer wing ribs 234. The outer wing ribs 234 may include outer wing transverse ribs 232 located at the inner end 230 of the outer wing. The outer wing transverse ribs 232 (… Figure 5 It can extend between the front wing spade 212 and the rear wing spade 214 of the outer wing, and can be arranged back-to-back with the center wing transverse rib 170. Figure 5 Mechanically connected.

[0113] Still referencing Figures 4 to 5 Similar to the arrangement of the center wing rib 170, the outer wing rib 232 can be formed of a metallic material to provide greater strength than the other outer wing ribs 234. The increased strength of the center wing rib 170 and the outer wing rib 232 allows the mechanically connected ribs to transfer loads between the outer wing structure 210 and the center wing structure 150. The outer wing forward spars 212 and the outer wing aft spars 214 can each extend continuously between the outer wing rib 232 and the wingtip 126, and can also be connected (e.g., via spars splices 216, 218) to the center wing forward spars 152 and the center wing aft spars 154 (i.e., at the wing joint 172) respectively for transferring loads between the outer wing structure 210 and the center wing structure 150.

[0114] The outer wing structure 210 may include an upper outer wing skin panel 240 and a lower outer wing skin panel 242, which are formed of composite materials and joined (e.g., bonded and / or mechanically fastened) to the outer wing front spars 212, the outer wing rear spars 214, and the outer wing rib 234, including the outer wing transverse rib 232. Similar to the skin panels of the aforementioned center wing structure 150, the area layer 302 constituting the upper outer wing skin panel 240 and the lower outer wing skin panel 242 may be a unidirectional layer 306. The area layer 302 may extend continuously from the outer wing front spars 212 to the outer wing rear spars 214 in the chord direction and continuously from the outer wing transverse rib 232 toward the wingtip 126 in the span direction. The panel thickness of the area layer 302 is 194 (…). Figure 20 The strength can be substantially constant, which simplifies the manufacture of the outer wing upper skin panel 240 and outer wing lower skin panel 242 by eliminating the need to incorporate ply reduction in area layer 302 to reduce bending stiffness along the outer direction.

[0115] exist Figures 4 to 5 In the middle, the panel thickness of the area layer 302 in the outer wing upper skin panel 240 and / or outer wing lower skin panel 242 is 194 ( Figure 20 The panel thickness 194 of the area layer 302 of the outer wing upper skin panel 190 and the outer wing lower skin panel 192 can be significantly thinner (e.g., at least 25%) than that of the area layer 302. Advantageously, the relatively thin specification (i.e., panel thickness 194) of the outer wing upper skin panel 240 and the outer wing lower skin panel 242 can improve the fatigue load performance (i.e., durability) of the skin panels relative to the fatigue load performance of aluminum skin panels of conventional wings. Furthermore, the relatively thin specification can minimize or avoid the possibility of interlaminar failure between the composite layers 300. The panel thickness 194 of the outer wing upper skin panel 240 and the outer wing lower skin panel 242 can be selected to provide the required level of stiffness and stability while resisting aerodynamic flutter. The minimum panel thickness 194 can be specified by the flammability requirements of the Federal Aviation Administration (FAA). For example, the upper skin panel and / or lower skin panel of the center wing structure 150 and the outer wing structure 210 may each have at least 32 composite layers 300, thereby producing a panel thickness 194 that provides a structural integrity level that meets FAA flammability requirements.

[0116] Similar to the aforementioned upper and lower center wing skin panels 190 and 192, in any of the outer wing structure embodiments disclosed herein, the area layers 302 of the upper and / or lower center wing skin panels 240 and 242 may include 0-degree layers, +45-degree layers, -45-degree layers, and 90-degree layers. The ratio of the 0-degree layer to the 90-degree layer may be between 1.75 and 2.5. For example, the upper and / or lower center wing skin panels 240 and 242 may include 40% 0-degree layers, 40% +45-degree and / or -45-degree layers, and 20% 90-degree layers. In another example, the upper and / or lower center wing skin panels 240 and 242 may include 38% 0-degree layers, 45% +45-degree and / or -45-degree layers, and 17% 90-degree layers.

[0117] refer to Figures 5 to 8 The center wing structure 150 may include one or more center wing stringers 270 for reinforcing the upper center wing skin panel 190 and / or the lower center wing skin panel 192. Each center wing stringer 270 may extend in the wingspan direction. For example, one or more center wing stringers 270 may extend continuously from the center wing box 122 and terminate at the center wing transverse rib 170. The center wing stringers 270 may be oriented generally parallel to each other (e.g., ±30 degrees) and may be spaced apart from each other in the chord direction. The center wing stringers 270 may be connected to the skin panels (upper center wing skin panel 190 and lower center wing skin panel 192) via stringer flanges 276. For example, the stringer flanges 276 may be adhesively (e.g., co-bonded, secondary-bonded, etc.) and / or mechanically fastened to the skin panels.

[0118] Still referencing Figures 5 to 8 Each center wing stringer 270 has at least one stringer flange 276 and at least one stringer web 278 extending outwardly from the stringer flange 276. As described below, each center wing stringer 270 may include a stringer runout 280 at the stringer end. Figure 20 As a means of gradually reducing the stiffness of the center wing stringer 270, the stringer web 278 can be gradually reduced in height to reduce the peeling force that would otherwise tend to separate the stringer flange 276 from the skin panel at the stringer end. In the example shown, the center wing stringer 270 is constructed as a blade stringer 274 with a single stringer web 278. However, the center wing stringer 270 can be provided in an alternative construction, such as a hat-shaped section stringer (not shown) with a pair of stringer webs extending outward from a pair of stringer flanges respectively, and the stringer webs can be interconnected by stringer caps.

[0119] like Figures 6 to 8As shown, the center wing rib 174 may include multiple rib cutouts 268 located in the outer flange 178 and / or the web 184, at discrete locations along the upper and / or lower edges of the center wing rib 174. The rib cutouts 268 provide clearance for the center wing stringers 270 to pass through the center wing rib 174. The center wing rib 174 may be connected to the upper skin panel 190 and / or the lower skin panel 192 of the center wing via a combination of coupling fasteners and mechanical fasteners 298. Additionally, the center wing rib 174 may be connected to the front wing spars 152 and the rear wing spars 154 of the center wing via mechanical fasteners 298.

[0120] refer to Figures 7 to 10 The forward spar 152 and the aft spar 154 of the center wing may each have a channel-shaped cross-section, which has a spar web 166 and a pair of opposing spar flanges 160 connected to each other through the spar web 166. Similar to the description below... Figure 13 The wing rib structure shown is as follows. Figures 9 to 10 The outer flange 160 of the spar can be connected to the web 166 of the spar via a radius portion 182. The web 166, the outer flange 160, and the radius portion 182 can be composed of a fabric layer 304. As described above, manufacturing the component from the fabric layer 304 can reduce or avoid warping and / or wrinkling, which could otherwise occur if the component were formed from a unidirectional layer 306.

[0121] refer to Figures 9 to 10 The outer flange 160 of the center wing forward spar 152 and / or center wing aft spar 154 may include one or more 0-degree layers to increase the shear strength of the outer flange 160. The 0-degree layer 308 can be described as a unidirectional layer 306, wherein the fibers are oriented along the longitudinal direction of the outer flange 160. In the illustrated example, the one or more 0-degree layers may be located approximately at the midplane of the outer flange 160. The 0-degree layers in the outer flange 160 may be excluded from the radial portion 182 and from the spar web 166 to avoid the aforementioned warping (i.e., spring-in) of the outer flange 160, which could otherwise occur during the curing of the unidirectional layer in the radial portion 182. During layup, the side edges of the 0-degree layers may be maintained at a distance of not less than 0.25 inches from the radial portion 182 to prevent warping during curing.

[0122] exist Figures 9 to 10 In the middle, the center wing forward sparb 152 and / or the center wing rear sparb 154 can be configured as a multi-piece structure to simplify the outer wing structure 210. Figure 5The manufacturing of the center wing. For example, the center wing forward sparb 152 and / or the center wing aft sparb 154 may each include a pair of spar chords 158, each spar chord having an L-shaped configuration including an inner sparb flange 162 and an outer sparb flange 160 interconnected by radial portions 182. The inner sparb flange 162 of each sparb chord 158 may be mechanically coupled to the sparb web 166, for example, via a double row of mechanical fasteners 298 extending through the inner sparb flange 162 and the sparb web 166. The outer sparb flange 160 of each sparb chord 158 may be fastened by adhesive (e.g., co-adhesive) and / or mechanical fastening, for example, via the double row of mechanical fasteners 298. Figures 7 to 8 ) is connected to the skin panel (e.g., the upper outer skin panel 190 of the center wing or the lower outer skin panel 192 of the center wing, Figures 7 to 8 As described above, the outer flange 160 of each sparb chord 158 may include one or more 0-degree layers to increase the shear strength of the outer flange 160. Although the figures show a multi-piece construction of the center wing forward sparb 152 and center wing aft sparb 154, such spars can be provided as a single-piece construction (not shown) consisting of a pair of outer flanges 160 interconnected by a sparb web 166, similar to... Figure 13 The following is the construction of the wing rib shown.

[0123] refer to Figure 4 and Figures 9 to 10 The center wing aft sparb 154 (and outer wing aft sparb 214) can be configured to handle higher bending loads than the center wing forward sparb 152 (and outer wing forward sparb 212). For example, the center wing aft sparb 154 (and outer wing aft sparb 214) may include a larger number of composite layers 300, resulting in a thicker cross-section and greater structural mass to provide correspondingly higher strength capabilities (e.g., lateral shear strength, bending strength) and higher bending stiffness than the center wing forward sparb 152 (and outer wing forward sparb 212). The center wing stringer 270 adjacent to the center wing aft sparb 154 may also be larger in mass than the center wing stringer 270 adjacent to the center wing forward sparb 152. The chordal difference in the strength capabilities of the center wing stringer 270 can be gradual or stepped. For example, the front 50% of the center wing stringer 270 (i.e., near the front wing spars 152 of the center wing) can have a single cross-sectional dimension, while the rear 50% of the center wing stringer 270 (i.e., near the rear wing spars 154 of the center wing) can have a smaller cross-sectional dimension.

[0124] refer to Figures 11 to 13 , Figure 11 An example of a center wing rib 174 connected to the center wing upper skin panel 190 and the center wing lower skin panel 192 is shown. Figure 12The central wing stringer 270, located on the upper skin panel 190 and the lower skin panel 192 of the central wing, is shown passing through the rib cutout 268 in the central wing rib 174. Figure 13 An example of a channel-shaped cross-section of the central wing rib 174 is shown, which may include opposing outer rib flanges 178 interconnected by a rib web 184. Each outer rib flange 178 may be connected to the rib web 184 via a radial portion 182. As described above, the central wing rib 174 may be made of a fabric layer 304 to reduce or avoid warping and / or wrinkling during curing.

[0125] The outer flange 178 may include one or more 0-degree layers located approximately at the midplane of the outer flange 178 to increase the shear strength of the outer flange 178. Similar to the center wing canard spar 152 ( Figure 9 ) and center wing rear spars 154 ( Figure 10 In the above arrangement, the 0-degree layer in the center wing rib 174 can be excluded from the radius portion 182 and the rib web 184 as a way to avoid or reduce warping (i.e., spring-in). Although shown as a single-piece construction, the center wing rib 174 can be provided as a multi-piece construction (not shown), wherein each rib flange 178 is part of a pair of rib chords 176 (e.g., upper and lower rib chords) interconnected by the rib web 184, similar to the center wing forespar 152. Figure 9 ) or center wing rear spars 154 ( Figure 10 (Multi-piece construction). Although not shown, the web 184 of the center wing rib 174 may each include discrete stiffeners to prevent out-of-plane buckling of the center wing rib 174.

[0126] refer to Figures 14 to 16 , Figure 14 The diagram shows a cross-sectional view of the center wing structure 150 at the center wing transverse rib 170. The upper center wing skin panel 190 and the lower center wing skin panel 192 are arranged in a manner similar to... Figures 6 to 8 The arrangement shown connects to the center wing transverse rib 170. Figure 15 A portion of the center wing cross rib 170, which is connected to the center wing front spar 152, is shown. Figure 16 A portion of the center wing transverse rib 170, which is connected to the center wing rear wing spars 154, is shown. As described above, the center wing transverse rib 170 and the outer wing transverse rib 232 ( Figure 5 It can have relative to the remaining center wing rib 174 and the remaining outer wing ribs 234 ( Figure 5The increased strength capacity is designed to accommodate the ends of the center wing stringer 270 and the outer wing stringer 272 at the center wing transverse rib 170 and the outer wing transverse rib 232, respectively. This is in addition to the spar splices (e.g., those connecting the outer wing structure 210 to the center wing structure 150) that form the spar splices. Figure 22-25 In addition to the load path provided by the center wing transverse rib 170 and the outer wing transverse rib 232, the load path can also be provided across the wing joint 172 from the outer wing structure 210 to the center wing structure 150.

[0127] Still referencing Figures 14 to 16 The center wing transverse rib 170 and the outer wing transverse rib 232 can further accommodate the difference in bending stiffness between the outer wing structure 210 and the center wing structure 150. For example, in an aircraft 100 with a relatively shallow, high aspect ratio wing and a short mean chord... Figure 1 In the case of ), the outer wing structure 210 ( Figure 1 The outer wing structure 210 can be moderately loaded and structurally sized based on dynamic loads, which are functions of mass and stiffness. The outer wing structure 210 can be generally flat and straight, and can have a smaller bending stiffness than the center wing structure 150 to allow for greater flexibility, thus meeting aeroelastic performance requirements. Conversely, the center wing structure 150 can be highly contoured and can include several high-load joints for attaching heavy components, including attaching the wing assembly 120 to the fuselage 102, and attaching the engine 106 (…). Figure 1 The landing gear (not shown) is mounted to the wing assembly 120 as described above.

[0128] refer to Figures 17 to 19 , Figures 17 to 18 An example of wing joint 172 is shown, illustrating its assembly configuration. Figure 17 ) and disassembly structure ( Figure 18 The center wing structure 150 includes a center wing transverse rib 170 and an outer wing transverse rib 232. As described above, the center wing structure 150 may include a center wing transverse rib 170 at the outer end 168 of each center wing. Each outer wing structure 210 may include an outer wing transverse rib 232 at the inner end 230 of the outer wing. At each wing joint 172, the outer wing transverse rib 232 and the center wing transverse rib 170 may be configured to be positioned back-to-back and mechanically fastened to each other. Figure 19 In one example, the center wing transverse rib 170 and the outer wing transverse rib 232 may each have a channel-shaped cross section, including a rib web 184 that connects a pair of rib flanges 178 to each other. However, the center wing transverse rib 170 and / or the outer wing transverse rib 232 may have alternative cross-sectional shapes.

[0129] like Figure 17As shown, when arranged back-to-back, the webs 184 of the center wing transverse rib 170 and the outer wing transverse rib 232 can be in direct physical contact with each other. The webs 184 of the outer wing transverse rib 232 and the center wing transverse rib 170 may include a mating hole pattern 296 for fastener holes. Figure 14 The fastener hole configuration is designed to receive a plurality of corresponding mechanical fasteners 298 (e.g., tension fasteners) for securing the outer wing transverse rib 232 and the center wing transverse rib 170 together. As described above, the center wing transverse rib 170 and the outer wing transverse rib 232 may have higher strength capabilities than the remaining wing ribs and may be composed of metallic materials such as steel or titanium or other high-strength alloys, which are compatible with the aforementioned composite materials of the spars, wing ribs, and stringers constituting the center wing structure 150 and the outer wing structure 210. However, the center wing transverse rib 170 and / or the outer wing transverse rib 232 may be formed primarily of fabric layers (not shown) to limit warping during curing.

[0130] refer to Figures 15 to 18 and Figure 20 The upper skin panel 190 and the lower skin panel 192 of the center wing can be included in the center wing transverse rib 170. Figures 17 to 18 The raised area 200 ( ) is located at the wing span direction and at the chord direction of the center wing sparsity 270. Figures 15 to 16 and Figure 20 Each padding area 200 consists of an area layer 302 that comprises the upper skin panel 190 and the lower skin panel 192 of the center wing. Figure 20 The composite layer 300 on top of ) Figure 20 The outer wing consists of an upper skin panel 240 and a lower skin panel 242, although not shown. These may also be included in the optional outer wing stringer 272 terminating at the outer wing transverse rib 232. Figures 17 to 18 A localized raised area 200 at the location of ). As described in more detail below, each raised area 200 causes the skin panel to bear a larger portion of the load, which would otherwise be borne by the stringers at the stringer transition section 280 ( Figure 20 The bearing capacity is reduced at the ) location, thereby reducing the peeling force on the edge of the stringer flange 276.

[0131] refer to Figure 20 An example is shown where the center wing stringer 270 terminates at the center wing transverse rib 170. As described above, each center wing stringer 270 has a stringer flange 276 and at least one stringer web 278 extending outward from the stringer flange 276. Additionally, each center wing stringer 270 has a stringer transition section 280 at its stringer end at the center wing transverse rib 170. Figure 20As shown, the stringer web 278 can gradually decrease in height along the stringer transition section 280 at its end. The gradual decrease in height of the stringer web 278 results in a gradual decrease in stringer stiffness, which advantageously reduces peeling forces that would otherwise tend to pull the stringer flange 276 away from the skin panel at the stringer end.

[0132] exist Figure 20 In this configuration, the panel thickness 194 of the lower center wing skin panel 192 (and the upper center wing skin panel 190) can gradually increase within the ramp region 202 towards the entire thickness of the raised region 200. The starting point of the ramp region 202 coincides with the starting point of the tapering portion of the stringer height at the stringer transition section 280. Within this ramp region, the panel thickness 194 can increase with a length-to-height ratio of not less than 2:1 to preferably provide a shallow ramp angle for the ramp region 202. The stringer flange 276 has a flange thickness 164 that can decrease with increasing panel thickness 194 within the ramp region 202. For example, the flange thickness 164 can decrease proportionally to the increase in panel thickness 194 within the ramp region, such that the combined thickness of the skin panel and the stringer flange 276 is constant within the ramp region 202.

[0133] The raised region 200 is preferably at least within the full thickness of the outer flange 178 of the central wing transverse rib 170. The increase in panel thickness 194 in the ramp region 202 and the corresponding decrease in flange thickness 164 of the stringer flange 276 facilitate bending loads at the wing joint 172 ( Figure 17 From the outer wing structure 210 () Figure 17 ) Transmitted to the central wing structure 150 ( Figure 17 Although not shown, the outer wing stringer 272 may also include a stringer transition section 280 at the outer wing transverse rib 232. The stringer transition section 280 of the outer wing stringer 272 may be similar to... Figure 20 The center wing stringer 270 is constructed in a manner that terminates at the end. Although not shown, the outer wing upper skin panel 240 and outer wing lower skin panel 242 may include raised areas 200 (and ramp areas 202) at locations complementary to the ends of the outer wing stringers 272. Such raised areas 200 in the outer wing upper skin panel 240 and outer wing lower skin panel 242 may be constructed in a manner similar to the aforementioned raised areas 200 in the center wing upper skin panel 190 and center wing lower skin panel 192.

[0134] Still referencing Figure 20Each central wing stringer 270's stringer flange 276 can be adhesively attached to either the upper central wing skin panel 190 (not shown) or the lower central wing skin panel 192. The adhesive interface between the stringer flange 276 and the skin panel can be described as an adhesive line 282. The central wing stringers 270 can be bonded to both the upper and lower central wing skin panels 190 and 192 by co-bonding uncured stringers to cured skin panels. Alternatively, bonding can include secondary bonding of cured stringers to cured skin panels via adhesive layer 284. In yet another example, bonding can include co-curing uncured stringers with uncured skin panels. In addition to bonding, the wing assembly 120 can include mechanical fasteners 298 to help secure the stringer flanges 276 to the skin panels.

[0135] refer to Figure 21 The adhesive line 282 (e.g., adhesive layer 284) between the lower skin panel 192 of the center wing and the stringer flange 276 of the center wing stringer 270 can terminate at a distance no greater than 0.50 inches from the stringer end. The termination of the adhesive line 282 at a distance spaced from the stringer end creates a flange-skin gap 286 between the stringer flange 276 and the skin panel. Within the flange-skin gap 286, the stringer flange 276 may not be bonded to the skin panel. The flange-skin gap 286 can be filled with a non-adhesive gap filler 288, such as a layer of fiberglass or other material that is not bonded to the stringer flange 276 and / or not bonded to the skin panel. By terminating the adhesive line 282 at a distance from the end of the stringer, the high shear stress 290 in the adhesive line 282 can be guided away from the stringer end due to wing bending. For example, Figure 21 The shear stress 290 and pull-out stress 292 in the bonding line 282 are shown. The relative magnitude of the pull-out stress 292 is indicated by the different lengths of the arrows. As shown, the maximum order of magnitude of the pull-out stress 292 is maintained at a certain distance from the end of the stringer. By maintaining a high pull-out stress 292 at a certain distance from the end of the stringer, the creep or shear stress 290 in the bonding line 282 between the stringer and the skin panel can be minimized, which can mitigate or eliminate the generation of cracks in the bonding line 282 that might otherwise propagate along the length of the center wing stringer 270.

[0136] refer to Figure 22The diagram shows a top view of the center wing lower skin panel 192 and the outer wing lower skin panel 242 at the wing joint 172. On one side of the wing joint 172, the center wing stringer 270 is mounted to the center wing lower skin panel 192 and terminates at the center wing transverse rib 170. On the opposite side of the wing joint 172, the outer wing stringer 272 is mounted to the outer wing lower skin panel 242 and terminates at the outer wing transverse rib 232. The stringer flanges 276 of the center wing stringer 270 and the outer wing stringer 272 can be opened at their ends as a way to improve the distribution of stringer loads to the respective skin panels, thereby reducing the adhesion line 282 between the stringer flange 276 and the skin panel. Figure 21 The peeling force in the wing joint 172. Although the outer wing stringer 272 is shown aligned with the center wing stringer 270, in an example not shown, the outer wing stringer 272 may not be aligned with the center wing stringer 270. In such an example, the padding area 200 on each side of the wing joint 172 ( Figure 20 It can provide a conduit for transferring loads from the outer wing structure 210 through the outer wing cross rib 232 and the center wing cross rib 170 to the center wing structure 150.

[0137] refer to Figures 23 to 26 An example of a splicing configuration for connecting the center wing rear spars 154 to the outer wing rear spars 214 at wing joint 172 is shown. The wing assembly 120 may include rear spars splice plates 218 that interconnect the spars ends 156 of the center wing rear spars 154 to the spars ends 156 of the outer wing rear spars 214 at each wing joint 172. Figures 25 to 26 As shown, the aft spar splice plate 218 can have a channel-shaped cross-section with a splice web 220 connecting a pair of opposing splice flanges 222. The channel-shaped cross-section of the aft spar splice plate 218 can be configured to nest within the channel-shaped cross-sections of the center wing aft spar 154 and the outer wing aft spar 214. The splice web 220 of the aft spar splice plate 218 can be horizontally oriented and can be positioned to correspond with the spar web 166 of both the center wing aft spar 154 and the outer wing aft spar 214. Figures 23 to 24 ) or inner flange of wing spars 162 ( Figure 25 The splicing flanges 222 on the upper and lower sides of the rear wing spars splice plate 218 can be positioned to contact the spars flanges 160 of the center wing rear wing spars 154 and the outer wing rear wing spars 214, and to contact the rib flanges 178 of the center wing transverse ribs 170 and the outer wing transverse ribs 232 on the upper and lower sides of the center wing rear wing spars 154 and the outer wing rear wing spars 214.

[0138] However, in other examples not shown, the aft spar splice 218 may only be connected to the spars web 166 of the center wing aft spar 154 and the outer wing aft spar 214. In yet another example, instead of being nested within the channel-shaped cross-section of the center wing aft spar 154 and the outer wing aft spar 214, the splice 218 may be mounted on the outer sides of the center wing aft spar 154 and the outer wing aft spar 214, which may be nested within the aft spar splice 218. Furthermore, instead of the splicing plate 218, which is an integral structure, the splicing plate 218 can have a multi-piece construction, including a web splicing plate (not shown) for the web plate 166 of the spar for connecting only the center wing rear spar 154 and the outer wing rear spar 214, a set of upper flange splicing strips (not shown) for connecting the upper spar flange of the center wing rear spar 154 and the outer wing rear spar 214, and a set of lower flange splicing strips (not shown) for connecting the lower spar flange of the center wing rear spar 154 and the outer wing rear spar 214.

[0139] exist Figures 23 to 25 In this example, two rows of mechanical fasteners 298 may extend through skin panels 190, 192, 240, 242, splice flange 222, and the outer flange 178 of the outer wing transverse rib 232 and the center wing transverse rib 170. Additionally, mechanical fasteners 298 may extend through splice web 220 and the spars web 166 of the outer wing rear spars 214 and the center wing rear spars 154. Although not shown, the wing assembly 120 may include a front spars splice plate 216 configured to connect the spars end 156 of the center wing front spars 152 to the spars end 156 of the outer wing front spars 212 in a manner similar to the aforementioned configuration of the rear spars splice plate 218, for connecting the center wing rear spars 154 to the outer wing rear spars 214. Advantageously, each wing joint 172 is configured to allow for on-site removal and replacement of the outer wing structure 210 from the center wing structure 150 by removing the front wing spars 216 and the rear wing spars 218 and disengaging the outer wing cross rib 232 from the center wing cross rib 170.

[0140] refer to Figure 27 In addition to one or more rows of mechanical fasteners 298 (e.g., two rows) that can extend along the length of each spar flange 160, the spar flanges 160 of the center wing forward spar 152 and center wing rear spar 154 can be bonded to the center wing upper skin panel 190 and center wing lower skin panel 192 via adhesive bonding along the length of the spar. Bonding the spar flanges 160 to the skin panels can include co-bonding the uncured spar to the cured skin panel. Alternatively, bonding can include secondary bonding of the cured spar to the cured skin panel via adhesive layer 284. Figure 20In yet another example, bonding may include co-curing uncured spars with uncured skin panels.

[0141] and Figure 21 Similar to the arrangement shown above, the adhesive line 282 (e.g., adhesive layer 284) between the skin panel and the outer flange 160 of the spar can terminate at a distance of no more than 0.50 inches from the spar end 156. The termination of the adhesive line 282 at a spaced distance from the spar end 156 creates a flange-skin gap 286 (e.g., ...). Figure 27 This gap can be filled with a non-adhesive gap filler 288 (e.g., glass fiber). By terminating each adhesive line 282 at a distance from the end of the spar terminal 156, the high shear stress 290 in the adhesive line 282 due to wing bending is reduced. Figure 21 The cracks can be guided away from the spar end 156, which can mitigate or eliminate the formation of cracks in the bonding line 282 that could otherwise propagate along the length of the spar outer flange 160.

[0142] Still referencing Figure 27 An example of a raised region 200 included in the lower skin panel 192 of the center wing is shown, located at the wing joint 172 for the spar terminal 156 of the rear wing spars 154 of the center wing. A similar raised region 200 may be included by the upper skin panel 190 of the center wing, the upper skin panel 240 of the outer wing, and the lower skin panel 242 of the outer wing for the spar terminals 156 of the front and rear wing spars of the center wing structure 150 and the outer wing structure 210. The raised region 200 for the spar terminal 156 may be constructed similarly to the raised region 200 described above for the center wing stringers 270 (and optionally the outer wing stringers 272). For example, each raised region 200 may include a ramp region 202 in which the panel thickness 194 of the skin panel locally increases along the spanwise direction toward the full thickness of the raised region 200. The outer flange 160 of the spar has a flange thickness 164, which can decrease as the panel thickness 194 within the ramp region 202 increases, so that the combined thickness of the skin panel and the outer flange 160 of the spar is constant within the ramp region 202.

[0143] Still referencing Figure 27Each raised region 200 is preferably at least within the full thickness of the corresponding outer flange 178 of the center wing transverse rib 170 and the outer wing transverse rib 232 (not shown). The increase in panel thickness 194 within the ramp region 202 and the corresponding decrease in flange thickness 164 of the spar outer flange 160 can facilitate the transfer of bending loads from the outer wing structure 210 to the center wing structure 150 at the wing joint 172. Each raised region 200 results in the skin panel bearing a larger portion of the load that would otherwise be borne by the spar at the spar end 156, thereby reducing the magnitude of the shear force in the adhesive layer 284, which could otherwise cause cracking in the adhesive layer 284, as described above.

[0144] refer to Figures 28 to 29 An example of an outer wing structure 210 is shown. As described above, the outer wing structure 210 includes a front wing spars 212 and a rear wing spars 214, a plurality of outer wing ribs 234 including outer wing transverse ribs 232, and an upper outer wing skin panel 240 and a lower outer wing skin panel 242. The aforementioned components of the outer wing structure 210 (i.e., spars, wing ribs, and skin panels) may define walls for one or more fuel tanks 128. In the example shown, the fuel tanks 128 in the outer wing structure 210 include a main tank 130 and a ventilation box 132 located outside the main tank 130.

[0145] As described above, the area layer 302 of the outer wing upper skin panel 240 and the outer wing lower skin panel 242 ( Figure 27 The panel thickness is 194 ( Figure 27 The thickness 194 of the outer wing upper skin panel 240 and outer wing lower skin panel 242 can be substantially constant, which simplifies manufacturing by eliminating the ply reduction 264 in area layer 302. Additionally, the panel thickness 194 of the outer wing upper skin panel 240 and outer wing lower skin panel 242 can be significantly thinner than the panel thickness 194 of the center wing upper skin panel 190 and / or center wing lower skin panel 192. The relatively thin specifications of the outer wing upper skin panel 240 and outer wing lower skin panel 242 can improve the fatigue load performance of the skin panels compared to conventional aluminum skin panels. As described in more detail below, the outer wing upper skin panel 240 and / or outer wing lower skin panel 242 can be composed of multiple outer wing stringers 272 and / or multiple 0-degree layer strips 252 interwoven within area layer 302 (e.g., Figures 40 to 43 )strengthen.

[0146] refer to Figures 30 to 32A cross-sectional view of an example of the outer wing structure 210 is shown, illustrating the outer wing rib 234 and multiple outer wing stringers 272 connecting to the upper outer wing skin panel 240 and the lower outer wing skin panel 242. The outer wing rib 234 extends between the outer wing front spar 212 and the outer wing rear spar 214 and may include rib cutouts 268 to allow the outer wing stringers 272, which extend from the outer wing transverse rib 232 along the wingspan, to pass through (…). Figures 28 to 29 Because the stiffness requirement of the outer wing portion outside the ventilation box 132 is reduced (due to the reduced inertial load outside the fuel tank 128), the outer wing stringer 272 may not be longer than the outer end 134 of the outermost fuel tank 128 (i.e., ventilation box 132). Figures 20 to 29 Extending further outward. However, in an example not shown, the outer wing structure 210 may be without the outer wing stringer 272, and the upper outer wing skin panel 240 and the lower outer wing skin panel 242 may each have a panel thickness 194 configured to provide the bending stiffness required for the outer wing structure 210. In another example, the upper outer wing skin panel 240 and / or the lower outer wing skin panel 242 may consist only of the 0-degree strip group 250, without the outer wing stringer 272.

[0147] refer to Figures 33 to 35 The diagram shows a chordal cross-sectional view of the outer wing structure, showing the outer end 134 of the outermost fuel tank 128 (e.g., ventilation box 132). Figures 28 to 29 The outer wing rib 234 is shown at the outer end 134 of the box. An optional outer wing stringer 272 is also shown, which may terminate at the outer wing rib 234 at the outer end of the box. Figures 34 to 35 Partial padding areas 200 for each outer wing spars 272 at the outer wing rib 234 at the outer end 134 of the box are shown. Partial padding areas 200 for the outer wing front spars 212 and outer wing rear spars 214 at the outer wing rib 234 are also shown.

[0148] refer to Figures 36 to 37 The diagram shows a spanwise cross-sectional view of a portion of the outer wing structure 210 at the outer end 134 of the box. As can be seen, the outer wing rib 234 at the outer end 134 of the box may have a larger cross-sectional area than the rest of the outer wing ribs 234 (e.g., thicker rib flange 178 and / or thicker rib web 184) to provide greater strength capacity, thereby helping to transfer bending loads between the outer wing rib 234 (e.g., at the outer end of the box) and the outer wing stringer 272.

[0149] like Figure 37 As shown, the lower outer wing skin panel 242 (and the upper outer wing skin panel 240) may include a partially raised area 200 at the outer wing rib 234 at the outer end 134 of the box for each stringer transition section 280. Figure 37 The raised area 200 shown can be related to the above. Figure 20 The raised area 200 for each center wing stringer 270 is constructed similarly. In this respect, the termination of the outer wing stringer 272 at the outer end 134 of the box can be the same as the termination of the center wing stringer 270 at the center wing transverse rib 170 described above. Figure 20 Similarly. For example, the stringer transition section 280 of each outer wing stringer 272 at the outer end 134 of the box may include a gradually decreasing stringer web 278 height as a way to gradually reduce the stiffness of the outer wing stringer 272, so as to reduce or avoid high peeling forces that might otherwise occur at the stringer end in the adhesive line 282 between the stringer flange 276 and the skin panel.

[0150] Additionally, the stringer flange 276 for each outer wing stringer 272 may include a gradually decreasing flange thickness 164 complementary to the gradually increasing panel thickness 194 in the ramp region 202. Similar to... Figure 21 The above arrangement is shown. Figure 37 The bonding line 282 (e.g., adhesive layer 284) between the outer wing spars 272 and the outer flange 160 can terminate at a distance (e.g., no more than 0.50 inches) from the end of the spars, thereby creating a flange-skin gap 286 that can be filled with non-adhesive gap filler 288 (e.g., fiberglass). As described above, the flange-skin gap 286 can minimize creep or shear stress 290 in the bonding line 282 between the outer wing spars 272 and the outer wing spars 272. Figure 21 This can mitigate or eliminate the formation of cracks in the bonding line 282, which could otherwise propagate along the length of the outer wing stringer 272.

[0151] refer to Figures 38 to 39 The diagram shows a cross-sectional view of a portion of the outer wing structure 210, which shows the outer wing forewing spars 212 at the outer wing rib 234 at the outer end 134 of the box. Figure 39 A raised region 200 and a ramp region 202 are shown in the lower skin panel 242 of the outer wing at the chord position of the outer wing front spar 212 at the outer end 134 of the box. The raised region 200 below the outer wing flange 160 can extend continuously (i.e., with a constant thickness) along the wingspan direction from the outer wing rib 234 at the outer end 134 of the box to the wingtip 126. A similar raised region 200 can be included in the upper skin panel 240 of the outer wing. The outer wing structure 210 can include raised regions 200 located in the upper skin panel 240 and the lower skin panel 242 of the outer wing for the outer wing rear spar 214 at the outer wing rib 234 at the outer end 134 of the box. The flange thickness 164 of the outer wing flange 160 can be similar to Figure 20 and Figure 37The arrangement shown above corresponds to a decrease in the thickness 194 of the skin panel within the ramp region 202 as the thickness increases. The bonding line 282 between the outer flange 160 of the spar and the skin panel can be continuous along the length of the outer wing's front spar 212 (and outer wing's rear spar 214).

[0152] refer to Figures 40 to 43 The diagram shows a cross-sectional view of the outer wing structure 210, illustrating examples of 0-degree stripe groups 250 incorporated into the upper skin panel 240 and / or lower skin panel 242 of the outer wing. Each 0-degree stripe group 250 extends along the wingspan direction. Each 0-degree stripe group 250 includes one or more 0-degree layer strips 252 interwoven within area layers 302. Each 0-degree layer strip 252 includes relatively narrow strips of 0-degree layer. The fibers in each 0-degree layer strip 252 are aligned with the wingspan direction. In some instances, one or more 0-degree stripe groups 250 can be derived from the outer wing transverse rib 232 ( Figure 29 )Continuously extends to the wingtip 126 ( Figure 29 In other instances, one or more 0-degree stripe groups 250 may extend from the outer wing transverse rib 232 to the outer end 134 of the box, or to a position beyond the outer end of the box in the wingspan direction. For example, as Figures 28 to 29 As shown, the two outermost wing pods 236 of the outer wing structure 210 may not have 0-degree strip groups 250 due to the relatively low stiffness requirements of the outermost wing pods 236.

[0153] refer to Figure 28 and Figure 40 Each 0-degree strip group 250 has a strip centerline 254 extending along the longitudinal direction of the 0-degree strip group 250. Figure 28 The 0-degree stripe groups 250 are spaced apart from each other such that the centerlines 254 of adjacent 0-degree stripe groups 250 are spaced 6-11 inches apart in the chord direction 256, and more preferably, 8-10 inches apart in the chord direction 256. The chord direction spacing 256 between the 0-degree stripe groups 250 of the skin panel is preferably uniform, as a way to maintain a uniform chord direction mass distribution of the outer wing structure 210.

[0154] refer to Figures 41 to 42 An example of a 0-degree strip group 250 incorporated into the lower skin panel 242 of the outer wing is shown. As described above, each 0-degree strip group 250 includes one or more 0-degree layer strips 252 interwoven within the area layer 302 of the skin panel. Figure 41An example of a 0-degree stripe group 250 with four 0-degree layer stripes 252 interwoven within area layer 302 is shown. However, the 0-degree stripe group 250 may include any number of 0-degree layer stripes 252. For example, the 0-degree stripe group 250 may include 1 to 70 0-degree layer stripes 252. The number of 0-degree layer stripes 252 in the 0-degree stripe group 250 may be specified by the stiffness requirements at the spanwise location.

[0155] To avoid or reduce the risk of layup cracking failure modes in the upper wing skin panel 240 or the lower wing skin panel 242, one or more 0-degree layer strips 252 of the 0-degree strip group 250 are preferably positioned between a pair of area layers 302. However, to prevent resin cracking, no more than four 0-degree layer strips 252 are stacked on top of each other. Although none of the 0-degree layer strips 252 defines the inner surface of the skin panel, the 0-degree layer strips 252 may be biased toward the inner surface of each skin panel.

[0156] refer to Figures 40 to 42 Each 0-degree layer strip 252 may have a strip width 258 ranging from 3 inches to 8 inches, and more preferably from 5 inches to 6 inches. However, the 0-degree layer strip 252 may have a strip width 258 less than 3 inches or greater than 8 inches. Figure 41 An example of a 0-degree strip group 250 is shown, wherein all 0-degree layer strips 252 have the same strip width 258, which simplifies manufacturing. As mentioned above, the 0-degree layer strips 252 can be offset toward the inner surface of the skin panel, and this can result in an effective wing thickness 262 between the area centroids of the 0-degree strip group 250. Figure 40 The effective wing thickness is increased compared to conventional wings that use traditional stringers (such as hat-shaped section stiffeners) to reinforce the skin panels.

[0157] Figure 42 An example of a 0-degree strip group 250 is shown, wherein the 0-degree layer strips 252 have different strip widths 258, the widest 0-degree layer strip 252 is located inside the skin panel (i.e., approximately in the mid-plane), and the narrower 0-degree layer strips 252 are located near the inner surface of the skin panel. Figure 42 The arrangement of the 0-degree stripe groups 250 can lead to improved strength or stiffness characteristics, while slightly increasing manufacturing complexity. Although not shown, the outer wing upper skin panel 240 may include one or more 0-degree stripe groups 250 configured similarly to the described arrangement of the 0-degree stripe groups 250 of the outer wing lower skin panel 242. As will be understood, the 0-degree stripe groups 250 can be provided in any of a variety of different cross-sectional configurations, geometries, and widths, and are not limited to these. Figures 41 to 42 The example shown.

[0158] refer to Figure 43 An example of a 0-degree strip group 250 in the lower skin panel 242 of the outer wing is shown. The 0-degree strip group 250 includes a ply reduction 264 of the 0-degree layer strips 252 to accommodate different bending stiffness requirements at different spanwise locations of the lower skin panel 242 of the outer wing. Although not shown, the upper skin panel 240 of the outer wing may also include one or more 0-degree strip groups 250 with ply reduction 264. Figure 43 As shown, the ply reduction 264 occurs only in the 0-degree layer strip 252, while the area layer 302 is continuous (without ply reduction), which simplifies manufacturing. Although Figure 43 The ply reduction 264 in the 0-degree layer strip 252 is shown, but the outer wing structure 210 may include one or more 0-degree strip groups 250, wherein there is no ply reduction 264 in the 0-degree layer strip 252.

[0159] The construction of the 0-degree strip group 250 in the outer wing upper skin panel 240 can differ from that in the outer wing lower skin panel 242 to accommodate different design standards and load conditions. For example, in the outer wing upper skin panel 240, the 0-degree strip group 250 can be interlaced to stabilize the outer wing upper skin panel 240 under maximum bending loads in a compressed state. Conversely, in the outer wing lower skin panel 242, the 0-degree strip group 250 can be interlaced to prevent the panel from cracking under tension during wing bending. As a result, the spanwise length, chordwise spacing 256, strip width 258, and / or number of strips of the 0-degree strip group 250 in the outer wing upper skin panel 240 can be greater than those of the 0-degree strip group 250 in the outer wing lower skin panel 242.

[0160] refer to Figures 44 to 45The diagram shows a cross-sectional view of an example of the outer wing structure 210 in the outer wing upper skin panel 240 and outer wing lower skin panel 242, each having outer wing stringers 272 mounted on a corresponding number of 0-degree strip groups 250. As described above, in one example, the outer wing stringers 272 can extend from the outer wing transverse rib 232 to the outer wing rib 234, thereby defining the outermost fuel tank 128's outermost end 134 in the outer wing structure 210. The 0-degree strip groups 250 can also extend from the outer wing transverse rib 232 and can extend towards the wingtip 126 beyond the outermost end 134. The outer wing stringers 272 can be more flexible in the spanwise direction than the center wing stringer 270. In this respect, the stringer web and / or stringer flange of the center wing stringer 270 may be larger or thicker than the stringer web and / or stringer flange of the outer wing stringer 272, so that the center wing structure 150 can accommodate the larger static load and higher stiffness requirements of the center wing structure 150 relative to the more moderately loaded and more flexible outer wing structure 210.

[0161] Figure 46 This is a flowchart of a method 500 for manufacturing wing assembly 120. Step 502 of method 500 includes attaching the inner end 230 of the outer wing of each of a pair of outer wing structures 210 to the opposite outer end 168 of the center wing of the center wing structure 150, respectively, to define a pair of wing joints 172 for attaching the outer wing structures 210 to the center wing structure 150. As described above, the center wing structure 150 has engine mounting positions on each opposite side of the wing centerline 124. At each wing joint 172, the sparsity ends 156 of the center wing forward spars 152 and the center wing aft spars 154 are located no more inward than the engine centerline 106 associated with the engine mounting positions, and no more than 10 percent of the distance between the engine centerline 106 and the wing centerline 124, further outward than the engine centerline 106.

[0162] Step 502, which connects each outer wing structure 210 to the center wing structure 150, may include connecting the outer wing transverse rib 232 on the inner end 230 of the outer wing of the outer wing structure 210 to the center wing transverse rib 170 on the outer end 168 of the center wing of the center wing structure 150, such that the outer wing transverse rib 232 and the center wing transverse rib 170 are arranged back-to-back. The rib webs 184 of the center wing transverse rib 170 and the outer wing transverse rib 232 may be positioned in direct physical contact with each other. Mechanical fasteners 298 may be installed in fastener holes in the rib webs 184 of the center wing transverse rib 170 and the outer wing transverse rib 232 to secure the rib webs 184 together.

[0163] Step 502, which connects each outer wing structure 210 to the center wing structure 150, may further include connecting the sparsity end 156 of the center wing front spars 152 to the sparsity end 156 of the outer wing front spars 212 via mechanical fasteners 298 using a front spars splice plate 216. Alternatively, method 500 may include connecting the sparsity end 156 of the center wing rear spars 154 to the sparsity end 156 of the outer wing rear spars 214 via mechanical fasteners 298 using a rear spars splice plate 218. As described above, each spars splice plate may have a channel-shaped cross-section that can be nested within the channel-shaped cross-sections of the front and rear spars. Mechanical fasteners 298 may be installed in fastener holes extending through the splice web 220 of each splice plate and through the sparsity web 166 of the center and outer wing spars. Mechanical fasteners 298 can also be installed in fastener holes extending through the skin panel, spar outer flange 160, and splice outer flange 222 to secure the outer spar to the center spar.

[0164] Method 500 may further include manufacturing a center wing structure 150, which may include laying a plurality of area layers 302 to form a center wing upper skin panel 190 and / or a center wing lower skin panel 192. As described above, the area layers 302 may include unidirectional layers 306, including 0-degree layers, +45-degree layers, -45-degree layers, and 90-degree layers, and no fabric layer 304. The area layers 302 may be laid such that the panel thickness 194 is substantially constant between the center wing forward spars 152 and the center wing aft spars 154 and along the wingspan direction at least from the fuselage 102 side to the center wing transverse rib 170. Method 500 may further include attaching the center wing upper skin panel 190 and the center wing lower skin panel 192 to the center wing forward spars 152, the center wing aft spars 154, and the center wing rib 174.

[0165] Method 500 may further include manufacturing each outer wing structure 210. Manufacturing each outer wing structure 210 may include laying multiple area layers 302 in a manner similar to laying the center wing upper skin panel 190 and the center wing lower skin panel 192 to form the outer wing upper skin panel 240 and / or the outer wing lower skin panel 242. For example, the area layers 302 may be laid such that the panel thickness 194 is constant between the outer wing front spars 212 and the outer wing rear spars 214 and along the wingspan direction from the outer wing transverse rib 232 to the wingtip 126.

[0166] Method 500 may further include interlacing one or more 0-degree layer strips 252 of one or more 0-degree strip groups 250 within the area layer 302 of the outer wing upper skin panel 240 and / or the outer wing lower skin panel 242 as a means of reinforcing the outer wing structure 210, as described above. After completing the outer wing structure 210, method 500 may include attaching the outer wing upper skin panel 240 and the outer wing lower skin panel 242 to the outer wing front spar 212, the outer wing rear spar 214, and the outer wing rib 234.

[0167] The step of laying the skin panel of the outer wing structure 210 or the center wing structure 150 may include laying one or more partial raised areas 200 on the inner surface of the skin panel. For example, in the center wing structure 150, the raised area 200 may be laid on the skin panel at the location of the center wing transverse rib 170 to accommodate the end of the center wing stringer 270 and to accommodate the center wing forward spars 152 and center wing aft spars 154 at the location of the center wing transverse rib 170. In each outer wing structure 210, the raised area 200 may be laid at the starting position of each outer wing stringer 272 at the outer wing transverse rib 232 and at the ending position of each outer wing stringer 272 at the outer wing rib 234 located at the outer end 134 of the box. The laying of any of the raised areas 200 may include a ramp area 202 on one or more sides of each raised area 200. In addition, the flange thickness 164 of the stringer flange 276 and / or the outer flange 160 of the wing beam can be gradually reduced in correspondence with the increase of the panel thickness 194 in the ramp region 202.

[0168] To minimize peeling forces in the adhesive line 282 that bonds the wing stringers to the skin panel, the method may include terminating the adhesive line 282 (e.g., adhesive layer 284) at a distance of no more than 0.50 inches from the stringer end in a manner that creates a stringer-skin gap between the stringer flange 276 and the skin panel. Method 500 may also include filling the flange-skin gap 286 with a non-adhesive gap filler 288 (e.g., a fiberglass layer) to prevent bonding between the stringer flange 276 and the skin panel at the stringer end. A similar process may be performed at the spar end 156 at the wing joint 172.

[0169] Method 500 may further include manufacturing the wing ribs of the center wing structure 150 and / or the outer wing structure 210 by laying the wing ribs into a channel-shaped cross-section having a rib web 184 and a pair of opposing rib flanges 178. As described above, the rib web 184 and rib flange 178 of each wing rib may consist of a stack of fabric layers 304. Method 500 may include weaving one or more 0-degree layers into the fabric layer 304 of the rib flange 178 to improve the shear strength of the rib flange 178. Method 500 may additionally include manufacturing the front spar and / or rear spar of the center wing structure 150 and / or the outer wing structure 210 in a similar manner.

[0170] The steps of manufacturing wing ribs may include manufacturing each wing rib as a multi-piece construction. For example, the method may include laying a pair of rib chords 176 (i.e., an upper rib chord and a lower rib chord), each rib chord having an L-shaped construction having an outer rib flange 178 and an inner rib flange 180 interconnected by a radial portion 182. The method may also include laying a generally planar rib web 184. Method 500 may then include bonding (e.g., co-bonding) and / or mechanically fastening the outer rib flanges 178 of the pair of rib chords 176 to the upper and lower skin panels of the center wing structure 150 or the outer wing structure 210, respectively. Method 500 may then include positioning the upper skin panel (and rib chord) relative to the lower skin panel (and rib chord) in an assembly fastener (not shown), and connecting the pair of rib chords 176 to each other via the rib web 184 by engaging (e.g., via mechanical fastener 298) the inner flange 180 of each rib chord to the rib web 184.

[0171] The manufacture of the front and / or rear spars of the center wing structure 150 and / or outer wing structure 210 may include manufacturing each sparb in a multi-piece construction by laying a pair of sparb chords 158 (i.e., upper and lower sparb chords), each sparb chord having an outer sparb flange 160 and an inner sparb flange 162 interconnected by a radial portion 182. Method 500 may include laying a generally planar sparb web 166. Additionally, the method may include bonding and / or mechanically fastening the outer sparb flanges 160 of the pair of sparb chords 158 to the upper and lower skin panels of the center wing structure 150 or outer wing structure 210, respectively. Method 500 may further include positioning the upper skin panel (and spar chord) relative to the lower skin panel (and spar chord) in an assembly fastener (not shown), and connecting the pair of spar chords 158 to each other via the spar web 166 by extending mechanical fasteners 298 through the inner flange 162 of each spar chord 158 into the spar web 166.

[0172] Figure 47This is a flowchart of a method 550 for mounting a wing assembly 120 having a pair of outer wing structures 210 connected to a center wing structure 150. Step 552 of the method includes providing a wing assembly 120 having a pair of outer wing structures 210, each outer wing structure being connected to the center wing structure 150. As described above, the center wing structure 150 includes a pair of outer wing tips 168 and a pair of engine mounting positions located on opposite sides of the wing centerline 124. Additionally, the center wing structure 150 includes a leading wing sparb 152 and a trailing wing sparb 154, each having a sparb end 156 at each outer wing tip 168. Each of the outer wing structures 210 includes an outer wing front sparb 212 and an outer wing rear sparb 214, respectively connected to the center wing front sparb 152 and the center wing rear sparb 154, to define a wing joint 172 connecting the outer wing structure 210 to the center wing structure 150. As described above, at each wing joint 172, the sparb ends 156 of the center wing front sparb 152 and the center wing rear sparb 154 are located no more inward than the engine centerline 106 associated with the engine mounting position, and no more than 10 percent of the distance between the engine centerline 106 and the wing centerline 124, further outward than the distance between the engine centerline 106 and the wing centerline 124.

[0173] Method 550 further includes step 554 of placing the wing assembly 120 under static ground load conditions, wherein the aircraft 100 may be substantially stationary (e.g., parked at an airport gate) and / or may be under static ground load conditions, wherein the wing assembly 120 is subjected to gravity due to the structural mass of the wing assembly 120. Method 550 further includes step 556 of placing the wing assembly 120 under dynamic load conditions, wherein the aircraft 100 may be in motion, and the wing assembly 120 is subjected to one or more dynamic loads. For example, a moving aircraft 100 may be subjected to gravity and inertial forces due to the structural mass of the wing assembly 120, aerodynamic loads on the wing assembly 120 during flight, and / or control surface loads on the wing assembly 120 due to flight deflection of one or more flight control surfaces (e.g., ailerons, flaps, and / or other control surfaces).

[0174] Now for reference Figures 48 to 58 An example of an outer wing structure 210 is shown, wherein the upper outer wing skin panel 240 and the lower outer wing skin panel 242 are reinforced by discrete slats 600, as an alternative to the aforementioned 0-degree strip group 250. Figures 54 to 55 As shown and described below, each slat 600 consists of a stack of slat layers 602, the width of which is narrower than the area layers 302 of the outer wing upper skin panel 240 and the outer wing lower skin panel 242. In addition to the slats 600, Figures 48 to 56 The structure of the outer wing structure 210 shown is the same as described above. Figures 28 to 39 The outer wing structure 210 shown is similar. For example, Figures 48 to 56 The outer wing structure 210 includes a front wing sparb 212 and a rear wing sparb 214, which are respectively connected to the front wing sparb 152 of the center wing. Figure 5 ) and center wing rear spars 154 ( Figure 5 ), to limit the connection of the outer wing structure 210 to the central wing structure 150 ( Figure 5 ) Wing joint 172 ( Figure 5 ).

[0175] Wing joint 172 may include front wing spars splice plate 216 ( Figure 5 ) and rear wing beam splice plate 218 ( Figure 5 The outer wing structure 210 includes a front wing spars 212 and 214 for connecting to the front wing spars 152 and 154 of the center wing structure 150, as described above. Additionally, the outer wing structure 210 includes an outer wing transverse rib 232 for connecting to the center wing transverse rib 170, as described above. The outer wing upper skin panel 240 and outer wing lower skin panel 242 each consist of an area layer 302 having a constant panel thickness along the wingspan direction, and are connected to the outer wing front spars 212, outer wing rear spars 214, and outer wing rib 234. As described above, the area layer 302 of the outer wing upper skin panel 240 and outer wing lower skin panel 242 includes a 0-degree layer and a 90-degree layer. In some examples, the ratio of the 0-degree layer to the 90-degree layer is between 1.75 and 2.5.

[0176] like Figures 50 to 51 As shown, multiple slats 600 are bonded to the inner surfaces of the outer wing upper skin panel 240 and the outer wing lower skin panel 242. The slats 600 extend along the wingspan. For example, each slat 600 can be extended from the outer wing transverse rib 232 ( Figures 48 to 49 ) Towards wingtip 126 ( Figures 48 to 49 ) Extension. The slat 600 can pass through one or more external wing nacelles 236 ( Figures 48 to 49 ) Continuous extension. In an example of an outer wing structure 210 with a fuel tank 128, the slats 600 may terminate at the outer end 134 of the tank. Figure 49 In this respect, the outer wing structure 210 may not have slats 600 on the outside of the fuel tank 128.

[0177] like Figures 48 to 50As shown, each slat 600 has a slat centerline 612 extending along the longitudinal direction of the slat 600. The slats 600 may be spaced apart from each other with a tangential spacing 256 of 6-11 inches between the slat centerlines 612. More preferably, the slats 600 may be spaced apart with a tangential spacing 256 of 8-10 inches. (This is in contrast to the 0-degree strip group 250...) Figure 40 Similar to the chordal spacing 256 between the slats 600, the chordal spacing 256 between the slats 600 can be uniform in order to maintain a substantially uniform chordal mass distribution in the outer wing structure 210. The slats 600 in the outer wing structure 210 can be respectively spaced with the center wing structure 150 (…). Figure 5 ) center wing stringer 270 ( Figure 5 Alignment. However, in other instances, the slats 600 in the outer wing structure 210 may not be aligned with the center wing stringer 270.

[0178] refer to Figure 51 Each of the slats 600 has a relatively low profile or a relatively small slat thickness 610. Figure 54 This results in the slats 600 and the associated skin panels (i.e., the slat-skin panel assembly 616) having a combined area centroid 260 that is closer to the inner surface of the skin panels 240, 242 than the combined area centroid of conventional stringers (e.g., hat-section stringers - not shown) and associated skin panels 240, 242. While the slat-skin area centroid 260 of the slat-skin panel assembly 616 is based on the combination of the cross-sectional areas of all the slats 600 (on the skin panels) with the cross-sectional areas of the skin panels 240, 242 between the front wing 212 and the rear wing 214, in Figure 51 For the purposes of this discussion, the slat-skin area centroid 260 of each slat-skin panel assembly 616 is represented by a slat-skin portion cross-sectional region 620, which comprises a combination of the cross-sectional region of a single slat 600 (i.e., the cross-shaded portion) and the cross-sectional region of the skin panel portions 614 of the skin panels 240, 242 extending outward from each side of the cross-shaded slat 600. The width of the skin panel portions 614 extends between midpoints 622 located on the cross-shaded slat 600 and between the slats 600 on the directly opposite sides of the cross-shaded slat 600. The proximity of the slat-skin area centroid 260 to the inner surfaces of the skin panels 240, 242 results in a larger effective wing thickness 262 than the effective wing thickness of a wing structure using conventional stringers to reinforce the skin panels. This larger effective wing thickness 262 translates into an increase in the bending load bearing efficiency of the outer wing structure 210. The increased bending load bearing efficiency corresponds to a reduction in the structural mass of the outer wing structure 210.

[0179] refer to Figure 51Each slat 600 can be configured to have a slat bending stiffness (i.e., in the spanwise direction) smaller than that of the combined slat-skin portion bending stiffness (i.e., in the spanwise direction) of the slat 600 and the skin panel portion 614 directly attached to and supported by the slat 600 (i.e., the upper skin panel 240 or the lower skin panel 242). For this purpose, each slat 600 can have a slat bending stiffness approximately 40-60% of the combined slat-skin portion bending stiffness. More preferably, the slat bending stiffness of each slat 600 is approximately 45-55% of the combined slat-skin portion bending stiffness. The slat bending stiffness can be the bending stiffness of the slat 600 at the spanwise location of the maximum slat thickness 610 (i.e., providing maximum slat stiffness). However, the slat bending stiffness of the slat 600 can be measured at any spanwise location along the slat 600. The combined slat-skin section bending stiffness is the combined stiffness of the slat 600 and the associated skin panel section 614, measured at the same spanwise location as the location where the slat bending stiffness is measured.

[0180] Maintaining the slat bending stiffness between 40-60% of the combined slat-skin section bending stiffness limits the magnitude of the peel force in the bond line (not shown) between the slats 600 and the skin panels 240, 242 to a level that prevents the peripheral edges of the slats 600 from separating from the inner surfaces of the skin panels 240, 242, while also preventing buckling of the unsupported sections of the skin panels 240, 242 between pairs of adjacent slats 600 when the outer wing structure 210 is under ultimate load (i.e., the maximum load expected during service). The chordal spacing 256 between the slats 600 ( Figure 51 The system can be positioned at a certain distance so that buckling of the skin panels 240 and 242 in the unsupported sections between adjacent strips 600 occurs at or above the ultimate load.

[0181] Still referencing Figure 51 As described above, each slat 600 may have a slat bending stiffness that is approximately 40-60% of the bending stiffness of the combined slat-skin portion. As described above, the cross-sectional area of ​​one of the slats 600 and the associated skin panel portion 614 is composed of… Figure 51 The intersecting shaded area of ​​the upper skin panel 240 and lower skin panel 242 is indicated and identified as the slat-skin section section region 620. For a slat 600 located between a pair of slats 600, the skin panel portion 614 has a panel portion width 618 extending between midpoints 622 located on opposite sides of the slats 600. Each midpoint 622 is located midway between the edges of adjacent slats 600, as shown in the image. Figure 51The upper skin panel 240 is shown. For a slat 600 located between a spar (e.g., the front spar 212 or the rear spar 214 of the outer wing) and another slat 600, the panel portion width 618 extends between the midpoint 622 between the side edge of the spar outer flange 160 on one side of the slat 600 and the edge of the adjacent slat 600 on the opposite side of the slat 600, as shown. Figure 51 The lower skin panel 242 is shown. In each instance, the panel portion width 618 of the skin panel portion 614 is centered on the slat 600.

[0182] The flexural stiffness of the skin panel portion 614 can be determined at least in part by the composition, number, and stacking order of the composite layers 300 constituting the skin panels 240 and 242, as well as by the slat-skin portion cross-sectional area 620, which can be the panel thickness 194, the panel portion width 618, and / or the chordal spacing 256. Figure 48 The function of ). Similarly, the magnitude of the slat bending stiffness of slat 600 can be determined at least in part by the layer composition, number of layers and stacking order of the slat layers 602 that make up slat 600, as well as the dimensions of slat 600 (e.g., slat width 604, slat thickness 610).

[0183] As described above, the slats 600, upper skin panel 240, and lower skin panel 242 can be configured to provide an outer wing structure 210 with a reduced bending stiffness compared to the increase in bending stiffness of the central wing structure 150. The increased flexibility of the outer wing structure 210 can accommodate the aeroelastic performance requirements of the outer wing structure 210. Furthermore, the increased flexibility of the outer wing structure 210 allows for a lighter wing assembly 120.

[0184] refer to Figures 52 to 53 The diagram shows a cross-sectional view of the outer wing structure 210, which illustrates the outer wing rib 234. The outer wing rib 234 may have rib cutouts 268 in the rib web 184 and / or rib flange 178 to provide clearance for the slats 600 to pass through. Alternatively, the rib flange 178 of the outer wing rib 234 may be partially shaped (not shown) to match the cross-sectional profile of each slat 600. As described above, the outer wing rib 234 may be mechanically fastened and / or coupled to the outer wing upper skin panel 240 and outer wing lower skin panel 242 and / or outer wing front spars 212 and outer wing rear spars 214.

[0185] refer to Figures 54 to 55The diagram shows a cross-sectional view of an example of slats 600 bonded to the lower skin panel 242 of the outer wing. As described below, each slat 600 can be laid and secondary bonded to the inner surface of the skin panels 240, 242. Alternatively, the slats 600 can be co-bonded or co-cured with the skin panels 240, 242. Each slat 600 can consist of 1-70 slat layers 602, more preferably, each slat 600 can include 10-70 slat layers 602, defining a slat thickness 610. The maximum number of slat layers 602 in a slat 600 can be determined by a 40-60% flexural stiffness ratio of the slat bending stiffness to the flexural stiffness of the combined slat-skin portion. The slat layers 602 can be unidirectional layers 306, which can include a combination of 0-degree layers and non-0-degree layers. Preferably, the 0-degree layer comprises the maximum percentage of slat layers 602 in each slat 600. Non-0 degree layers may include 90 degree layers, +45 degree layers, -45 degree layers, or other layer orientations. 0 degree layers have fibers oriented in a longitudinal direction parallel to the slat 600.

[0186] Each slat 600 may have a slat width 604 between 3 and 6 inches. More preferably, each slat 600 may have a slat width 604 between 4 and 5 inches. However, slats 600 may have a slat width 604 less than 3 inches or greater than 6 inches. Figure 54 In the example, the slat layers 602 within the slat 600 have the same slat width 604, resulting in a rectangular cross-sectional shape for the slat 600. However, the slat layers 602 within the slat 600 can have different slat widths 604, and this can result in a non-rectangular cross-sectional shape. For example, the slat 600 can have a trapezoidal cross-sectional shape, such as... Figure 55 As shown. In this arrangement, each slat 600 may have a relatively tapered side edge 606 that gradually decreases with an edge taper angle 608. In some instances, the edge taper angle 608 may be no greater than 30 degrees relative to the inner surface of the skin panels 240, 242.

[0187] refer to Figure 56 The diagram shows a cross-sectional view of the outer wing structure 210 along its span, illustrating an example of ply reduction 264 in the slat layer 602 of the slats 600. As described above, each slat 600 has a slat thickness 610 ( Figure 54In the illustrated example, the slat thickness 610 of one or more slats 600 in the outer wing structure 210 may gradually decrease in the spanwise direction as a way to gradually reduce slat stiffness in the spanwise direction. For this purpose, the slats 600 in the outer wing structure 210 may include ply reductions 264 in one or more slat layers 602. The spanwise location of the ply reductions 264 can be selected based on the desired bending stiffness distribution of the skin panels 240, 242 along the spanwise direction and can be based on predicted aerodynamic wing loads on the outer wing structure 210 during flight. Alternatively, the outer wing structure 210 may be configured such that the slats 600 have a constant slat thickness 610 along the longitudinal direction of the slats 600. Preferably, the slats 600 within each outer wing structure 210 have the same construction (e.g., the same width, the same cross-sectional shape, the same layer stack, and / or the same material composition) to simplify manufacturing. In addition, each slat 600 within the outer wing structure 210 preferably has the same construction as a way to keep the chordal mass distribution of the slats 600 within 10% of each other.

[0188] Figure 57 It is to manufacture products with the features described above. Figures 48 to 56 A flowchart of a method 700 for a wing assembly 120 of an outer wing structure 210 is shown. Method 700 includes a step 702 of bonding one or more slats 600 to the inner surface of at least one of the outer wing upper skin panel 240 and outer wing lower skin panel 242 of the outer wing structure 210. As described above, each slat 600 extends in the wingspan direction and is composed of a stack of slat layers 602. Also as described above, the outer wing upper skin panel 240 and outer wing lower skin panel 242 are each composed of an area layer 302 having a constant panel thickness along the wingspan direction. The slat layer 602 has a width narrower than the area layer 302.

[0189] Manufacturing with Figures 48 to 56 The method 700 for manufacturing the wing assembly 120 of the outer wing structure 210 shown above may include manufacturing a wing assembly 120 having, as shown above, the wing assembly 120 of the outer wing structure 210 shown above. Figures 28 to 45 Any one or more of the above steps of the method 500 for the wing assembly 120 of the outer wing structure 210 shown, except as... Figures 41 to 45 The method 700, which involves weaving the 0-degree strip group 250 into the skin panels 240, 242 of the outer wing structure 210, includes, as shown, the following: Figures 48 to 56 The slats 600 are shown attached to the skin panels 240 and 242 of the outer wing structure 210.

[0190] Before bonding the slats 600 to the skin panels 240, 242, method 700 may include laying a plurality of 0-degree layers and non-0-degree layers to produce the slats 600. The 0-degree layers may include at least one slat layer 602 representing the largest percentage of the slats 600. As described above, each slat 600 may include between 1 and 70 slat layers 602, more preferably between 10 and 70 slat layers 602. Method 700 may include laying the slat layers 602 such that at least some of the slats 600 have a slat width 604 between 3 and 6 inches, more preferably between 4 and 5 inches. Method 700 may also include laying the slat layers 602 such that at least some of the slats 600 have a slat width 604 such as… Figure 54 The rectangular cross-section shown is as described above.

[0191] However, in other instances, method 700 may include laying slat layers 602 such that at least some slats 600 have opposing slat side edges 606 oriented with an edge taper angle 608 of no more than 30 degrees relative to the inner surface of the respective upper skin panel 240 or lower skin panel 242, and resulting in... Figure 55 The slats 600 shown have a trapezoidal cross-section. More preferably, the slat layer 602 can be laid such that each slat side edge 606 forms an edge cone angle 608 between 15 and 30 degrees, so as to minimize stress concentration in the skin panels 240, 242 due to the slat side edges 606.

[0192] Method 700 may include positioning slats 600 on each of the outer wing upper skin panel 240 and lower skin panel 242 with a chordal spacing 256 of 6-11 inches between slat centerlines 612. As described above, each slat 600 has a slat centerline 612 extending along the longitudinal direction of the slat 600. In some instances, slats 600 may be positioned on the outer wing upper skin panel 240 and outer wing lower skin panel 242 such that when the outer wing structure 210 is attached to the center wing structure 150, the slats 600 are aligned with the center wing stringer 270 of the center wing structure 150, respectively. However, slats 600 may be positioned such that the slats 600 are not aligned with the center wing stringer 270.

[0193] Method 700 may further include laying each slat 600 to have approximately 40-60% of the combined slat-skin portion flexural stiffness of the slat 600 and the skin panel portion, as described above regarding... Figure 51 Described. More preferably, the slats 600 can be laid to have approximately 45-55% of the slat bending stiffness for the composite slat-skin portion. Method 700 may also include at least one ply reduction 264 in one or more slat layers 602, such as Figure 56As shown. As described above, including one or more ply reductions 264 in the slat layer 602 can gradually reduce the slat bending stiffness along the spanwise direction. The spanwise position of the ply reductions 264 can be selected to produce the desired bending stiffness distribution of the skin panels 240, 242 along the spanwise direction, and can be based on the predicted aerodynamic wing loads on the outer wing structure 210 during flight.

[0194] Method 700 may further include laying the upper wing skin panel 240 and associated slats 600 in a manner having a combined axial stiffness in the spanwise direction, which is higher than the combined axial stiffness of the lower wing skin panel 242 and associated slats 600. The higher axial stiffness of the upper skin panel 240 can accommodate relatively higher compressive loads in the upper skin panel 240 during upward bending of the outer wing structure 210. Conversely, due to the high tensile strength of the fibers, the composite structure has higher tensile strength, so the lower skin panel 242 may require a lower axial stiffness.

[0195] Method 700 may further include laying slats 600 to a length such that, when coupled to the outer wing upper skin panel 240 or the outer wing lower skin panel 242, each slat 600 terminates at a position no further outward than the outer end 134 of the fuel tank 128 that may be included in the outer wing structure 210. As described above, the slats 600 may extend continuously from the outer wing transverse rib 232 toward the wingtip 126. As described above, due to the reduced inertial load at the location outside the fuel tank 128 (i.e., due to the reduced mass), the stiffness requirement in the outer wing nacelle 236 outside the fuel tank 128 can be relatively lower compared to the bending stiffness requirement inside the outer end 134.

[0196] The step of bonding the slats 600 to the inner surfaces of the skin panels 240, 242 may include individually laying and curing each slat 600, and then bonding each slat 600 to the skin panels 240, 242 via an adhesive layer (not shown). Alternatively, the slats 600 (in a cured or uncured state) may be laid individually and co-bonded with the skin panels 240, 242 in their uncured or cured states. In another example, the slats 600 may be laid individually and co-cured with the skin panels 240, 242 in a single curing cycle.

[0197] In addition to laying the slat layer 602 to form the slat 600, method 700 may include providing a plurality of outer wing ribs 234 with rib cutouts 268 to provide gaps for the slat 600 to pass through the outer wing ribs 234. As described above, the outer wing ribs 234 (of the composite layer 300) may be laid in a channel-shaped cross section. The method may include connecting the outer wing ribs 234 to the outer wing front spar 212 and the outer wing rear spar 214.

[0198] Step 704 of method 700 includes attaching the upper outer wing skin panel 240 and the lower outer wing skin panel 242 to the outer wing front spar 212, the outer wing rear spar 214, and the outer wing rib 234, which includes the outer wing structure 210 transverse rib 232. As described above, the upper outer wing skin panel 240 and the lower outer wing skin panel 242 can be adhesively and / or mechanically fastened to the outer wing front spar, the outer wing rear spar, and the outer wing rib 234.

[0199] Step 706 of method 700 is substantially similar to step 502 of method 500 described above. Step 706 includes attaching the inner end 230 of the outer wing of each of a pair of outer wing structures 210 to the opposite outer end 168 of the center wing of the center wing structure 150, respectively, to define a pair of wing joints 172. As described above and Figure 4 As shown, the center wing structure 150 has engine mounting positions on each opposite side of the wing centerline 124. At each wing joint 172, the spars ends 156 of the center wing forward spars 152 and center wing aft spars 154 are located no more inward than the engine centerline 106 associated with the engine mounting positions, and no more than 10 percent of the distance between the engine centerline 106 and the wing centerline 124, further outward from the engine centerline 106.

[0200] Additionally, step 706 may include any one or more of the aforementioned alternative or additional steps related to step 502. For example, step 706 may include connecting the outer wing transverse rib 232 on the inner end 230 of the outer wing structure 210 to the center wing transverse rib 170 on the outer end 168 of the center wing structure 150, such that the outer wing transverse rib 232 and the center wing transverse rib 170 are positioned as follows: Figure 17 The back-to-back arrangement is shown and described above. Step 706 may further include using a front spar splice plate 216 to mechanically fasten the spar terminal 156 of the center wing front spar 152 to the spar terminal 156 of the outer wing front spar 212, and using a rear spar splice plate 218 to mechanically fasten the spar terminal 156 of the center wing rear spar 154 to the spar terminal 156 of the outer wing rear spar 214, as shown. Figure 5 As shown and described above.

[0201] Figure 58 It is loaded with the features described above Figures 48 to 56A flowchart of a method 800 for a composite wing assembly 120 of an outer wing structure 210 is shown. The method includes step 802 of providing a wing assembly 120 having a pair of outer wing structures 210, each outer wing structure being connected to a central wing structure 150. As described above, the central wing structure 150 includes a pair of outer wing tips 168 and a pair of engine mounting positions located on opposite sides of the wing centerline 124. Additionally, the central wing structure 150 includes a front wing sparb 152 and a rear wing sparb 154, each having a sparb end 156 at each outer wing tip 168. Figures 48 to 56 The outer wing structure 210 includes a front wing sparb 212 and a rear wing sparb 214, which are configured to be connected to the front wing sparb 152 and the rear wing sparb 154 of the center wing, respectively, to define a wing joint 172 connecting the outer wing structure 210 to the center wing structure 150. Additionally, the outer wing structure 210 includes an upper wing skin panel 240 and a lower wing skin panel 242, each consisting of an area layer 302 having a constant panel thickness along the wingspan direction, and connected to the front wing sparb 212, the rear wing sparb 214, the upper wing skin panel 240, and the lower wing skin panel 242. As described above, the upper wing skin panel 240 and / or the lower wing skin panel 242 include one or more strips 600 bonded to the inner surfaces of the skin panels 240, 242. Each slat 600 extends in the spanwise direction and consists of a stack of slat layers 602 that are narrower than the area layers 302 of the skin panels 240, 242. As described above, the center wing structure 150 is configured such that the spar ends 156 of the center wing forward spar 152 and the center wing aft spar 154 are located at each wing joint 172 no more inward than the engine centerline 106 associated with the engine mounting position, and no more outward than 10 percent of the distance between the engine centerline 106 and the wing centerline 124.

[0202] Step 804 of method 800 includes placing the wing assembly 120 under ground static load conditions in a manner similar to that in step 554 of method 550 described above. Step 806 of method 800 includes placing the wing assembly 120 under dynamic load conditions. As described above with respect to method 550, the ground static load conditions relate to the aircraft 100 being substantially stationary and under ground static load, under which the wing assembly 120 is subjected to gravity due to the structural mass of the wing assembly 120. The dynamic load conditions relate to the motion of the aircraft 100 and the wing assembly 120 being subjected to one of the following: gravity due to the structural mass of the wing assembly 120, inertial forces due to the structural mass of the wing assembly 120, aerodynamic loads on the wing assembly 120, and / or control surface loads on the wing assembly 120.

[0203] Now for reference Figures 59 to 72 An example of an outer wing structure 210 is shown, wherein the upper outer wing skin panel 240 and the lower outer wing skin panel 242 are reinforced by outer wing stringers 272, as... Figures 48 to 56 The above-mentioned slat 600 is an alternative, and as... Figures 28 to 43 The above-mentioned alternative to the 0-degree strip group 250. For example... Figures 61 to 62 As shown and described below, the outer wing structure 210 includes one or more outer wing stringers 272 connected to the upper outer wing skin panel 240 and / or the lower outer wing skin panel 242. Each outer wing stringer 272 extends in the spanwise direction and has a cap-shaped section 900. Each outer wing stringer 272 has a pair of stringer flanges 276, a pair of stringer webs 278, and a stringer cap 910. The stringer cap 910 includes one or more 0-degree layer strips 252 interwoven within the stringer layer 908 of the stringer cap 910. Figures 65 to 66 In addition to the outer wing stringer 272, Figures 59 to 70 The outer wing structure 210 shown is similar in construction to the one described above. Figures 48 to 56 The outer wing structure 210 and the above Figures 28 to 43 The outer wing structure 210.

[0204] Figures 59 to 62 Multiple outer wing stringers 272 are shown bonded to the inner surfaces of the outer wing upper skin panel 240 and the outer wing lower skin panel 242. Each outer wing stringer 272 extends continuously from the outer wing transverse rib 232 toward the wingtip 126 along the wingspan. For an outer wing structure 210 including a fuel tank 128, the outer wing stringer 272 may terminate at the outer end 134 of the tank. However, in other embodiments, one or more outer wing stringers 272 may extend continuously from the outer wing transverse rib 232, across the fuel tank 128, to the wingtip 126, spanning multiple outer wing nacelles 236.

[0205] like Figure 59 and Figure 61 As shown, each outer wing stringer 272 has a stringer centerline 912 extending along the longitudinal direction of the outer wing stringer 272. The outer wing stringers 272 may be spaced apart from each other by a chordal spacing 256 of 6-11 inches between the stringer centerlines 912. As described above, the chordal spacing 256 of the outer wing stringers 272 may be uniform with respect to the chordal spacing 256 of the 0-degree strip group 250 and the chordal spacing 256 of the slats 600, in order to maintain a uniform chordal mass distribution in the outer wing structure 210. The outer wing stringers 272 may optionally be connected to the center wing structure 150 (… Figure 5 ) center wing stringer 270 ( Figure 5 Alignment. Alternatively, the outer wing stringer 272 can be positioned out of alignment with the center wing stringer 270.

[0206] refer to Figure 62 The figure shows a portion of the outer wing structure 210, illustrating an example of an outer wing stringer 272 connecting the upper outer wing skin panel 240 and the lower outer wing skin panel 242. As shown, the outer wing stringer 272 has a relatively low stringer height 914, which results in its area centroid being closer to the area centroid 260 of the skin panels 240, 242 than that of a conventional, larger hat-shaped stringer. The low-profile outer wing stringer 272 and the associated skin panels (i.e., the stringer-skin panel assembly) have a combined area centroid 260 that is closer to the inner surface of the skin panels 240, 242 than the combined area centroid of a conventional stringer and associated skin panel. Although the stringer-skin panel assembly's stringer-skin area centroid 260 is based on the combination of the cross-sectional areas of all outer wing stringers 272 (on the skin panel) and the cross-sectional areas of the skin panels 240, 242 between the front wing spars 212 and the rear wing spars 214, for the purposes of this discussion, Figure 62 In this configuration, the centroid 260 of the stringer-skin area of ​​each stringer-skin panel assembly is represented by a stringer-skin portion cross-sectional region 920, which comprises a combination of the cross-sectional region of a single outer wing stringer 272 (i.e., the cross-shaded portion) and the cross-sectional region of the skin panel portions 614 of the skin panels 240, 242 extending from each side of the cross-shaded outer wing stringer 272. The width of the cross-shaded skin panel portion 614 extends between the midpoints 622 of the outer wing stringers 272 located on the directly opposite sides of the cross-shaded outer wing stringers 272. The proximity of the centroid 260 of the stringer-skin area to the inner surfaces of the skin panels 240, 242 results in a greater effective wing thickness 262 than that of wings using conventional stringers to reinforce the skin panels. The greater effective wing thickness 262 translates into an increase in the bending load bearing efficiency of the outer wing structure 210, and this can lead to a reduction in the structural mass of the wing assembly 120.

[0207] Still referencing Figure 62 Similar to the above and Figures 48 to 58 The bending stiffness ratio related to the slats 600, the outer wing stringer 272 can have a stringer bending stiffness lower than that of the combined stringer-skin portion bending stiffness of the outer wing stringer 272 and the skin panel portion 614 directly supported by the outer wing stringer 272. In this respect, Figures 59 to 70 Each outer wing stringer 272 may have a stringer bending stiffness that is approximately 40-60% (more preferably, 45-55%) of the combined stringer-skin portion bending stiffness as described above. The stringer bending stiffness can be the bending stiffness of the outer wing stringer 272 at the spanwise location where the maximum stringer bending stiffness is achieved. However, the stringer bending stiffness of the outer wing stringer 272 can be at any spanwise location along the outer wing stringer 272. The combined stringer-skin portion bending stiffness is the combined stiffness of the outer wing stringer 272 and the associated skin panel portion 614 measured at the same spanwise location as where the stringer bending stiffness is measured.

[0208] As mentioned above Figures 28 to 43 210 outer wing structure and Figures 48 to 56 As described in the outer wing structure 210, maintaining the stringer bending stiffness of each outer wing stringer 272 between 40-60% of the bending stiffness of the combined stringer-skin portion limits the magnitude of the peel force in the bond line between the outer wing stringers 272 and the skin panels 240, 242 to a level that prevents the peripheral edges of the outer wing stringers 272 from separating from the inner surfaces of the skin panels 240, 242, while also preventing buckling of the unsupported sections of the skin panels 240, 242 between pairs of adjacent outer wing stringers 272 when the outer wing structure 210 is under ultimate load. The chordal spacing 256 between the outer wing stringers 272... Figure 59 This can be such that buckling in the unsupported sections of the skin panels 240, 242 between adjacent outer wing stringers 272 occurs at or above the ultimate load.

[0209] Still referencing Figure 62 As described above, the cross-sectional area of ​​an outer wing stringer 272 and the associated skin panel portion 614 is represented by the aforementioned stringer-skin portion cross-sectional area 920. For those located as shown in... Figure 62 The outer wing stringers 272 in the upper skin panel 240 are located between one pair of outer wing stringers 272. The skin panel portion 614 has a panel portion width 618 extending between midpoints 622 located on opposite sides of the outer wing stringers 272. As described above, each midpoint 622 is located midway between the edges of adjacent outer wing stringers 272. For the spars 212, 214 and as shown in the figure, Figure 62The outer wing stringer 272 in the lower skin panel 242 has a panel portion width 618 that extends between the outer wing stringers 272 and the midpoint 622 between the outer wing stringer flange 160 of the outer wing front spar 212 or outer wing rear spar 214 on one side of the outer wing stringer 272 and the edge of the adjacent outer wing stringer 272 on the opposite side of the outer wing stringer 272. In each instance, the panel portion width 618 of each skin panel portion 614 is centered on the outer wing stringer 272.

[0210] As mentioned above Figures 28 to 43 and Figures 48 to 56 The construction of the outer wing stringer 272 is described above. Figure 62 The flexural stiffness of the skin panel portion 614 can depend on the layer composition, number of layers, and stacking order of the composite layer 300 of the skin panels 240 and 242, and on the cross-sectional area of ​​the stringer-skin portion, which can be the panel thickness 194, the panel portion width 618, and / or the chordal spacing 256 between the outer wing stringers 272. Figure 61 The bending stiffness of the outer wing stringer 272 can be a function of the stringer structure 908 comprising the outer wing stringer 272. Similarly, the bending stiffness of the stringer can be determined by the layer composition, number of layers, and stacking order of the stringer layers 908 that make up the outer wing stringer 272, and can also be determined by the dimensions of the outer wing stringer 272 (e.g., stringer width 918, stringer height 914—). Figure 65 ).

[0211] refer to Figures 63 to 64 The diagram shows a cross-sectional view of the outer wing structure 210, illustrating an outer wing rib 234 extending between the front wing spars 212 and the rear wing spars 214. One or more outer wing ribs 234 may have one or more rib cutouts 268 in the rib web 184 and / or rib flange 178 to provide clearance for the outer wing stringers 272 through the outer wing ribs 234. Alternatively, the rib flanges 178 of the outer wing ribs 234 may be partially shaped to match the profile of each outer wing stringer 272.

[0212] refer to Figures 65 to 66 , Figure 65A cross-sectional view is shown of an example of an outer wing stringer 272 bonded to an outer wing lower skin panel 242. The outer wing stringer 272 can be laid out and secondary bonded to the inner surfaces of skin panels 240, 242. Alternatively, the outer wing stringer 272 can be co-bonded or co-cured with skin panels 240, 242. As described above, each outer wing stringer 272 may include a pair of stringer flanges 276, a pair of stringer webs 278, and stringer caps 910 interconnected to the stringer webs 278. Each stringer flange 276 can be interconnected to the stringer webs 278 via a radius portion 182. Similarly, each stringer web 278 can be interconnected to the stringer via a radius portion 182. As described below, the stringer flanges 276, stringer webs 278, and stringer caps 910 may be composed of a stringer layer 908. The stringer layer 908 may be a fabric layer 304. The outer wing stringer 272 can exclude the use of unidirectional layer 306 as a way to avoid warping, which would otherwise be associated with unidirectional layer 306 in a composite structure with a non-planar shape.

[0213] Still referencing Figures 65 to 66 To increase the bending stiffness of the outer wing stringers 272, the stringer cover 910 includes one or more 0-degree layer strips 252 interwoven within the stringer layer 908. Each 0-degree layer strip 252 in the stringer cover 910 can extend across the entire width of the stringer cover 910. However, to prevent warping of the outer wing stringers 272 during curing, each 0-degree layer strip 252 is confined to the stringer cover 910 and excluded from the radial portion 182 on either side of the stringer cover 910. Although Figure 65 An example of a stringer cap 910 with three 0-degree layer strips 252 is shown, but the stringer cap 910 may include any number of 0-degree layer strips 252. For example, the stringer cap 910 may include 1-10 or more 0-degree layer strips 252, which may be necessary to meet flexural stiffness requirements. To prevent resin cracking, no more than four 0-degree layer strips 252 may be directly stacked on top of each other in the stringer cap 910.

[0214] Each outer wing stringer 272 may have a stringer width 918 of up to 4 inches. The stringer width 918 may be measured on each side of the stringer vertical axis 916 between the intersection of the stringer flange 276 and the stringer web 278. Each outer wing stringer 272 may have a stringer height 914 of no more than 3 inches, which may advantageously result in the area centroid 260 being close to the inner surface of the skin panels 240, 242, and may result in a relatively large effective wing thickness 262, as described above. Figure 62As shown. In some examples of the outer wing stringer 272, the stringer flanges 276 may each have a flange-side edge 906 that gradually decreases in angle relative to the inner surface of the corresponding upper or lower skin panel 240 or skin panel 242 bonded to the outer wing stringer 272 with an edge taper angle 608 not exceeding 30 degrees. Preferably, the edge taper angle 608 is between approximately 15 and 30 degrees to reduce or prevent indentations (i.e., dents) in the skin panels 240, 242.

[0215] refer to Figure 65 and Figure 67 The outer wing stringer 272 may include a stringer base laminate 902. The stringer base laminate 902 may consist of unidirectional layers 306 and may be positioned between the stringer flange 276 of each outer wing stringer 272 and the upper or lower outer wing skin panel 240. The stringer base laminate 902 may include one or more 0-degree layer strips 252. To prevent resin cracking, no more than four 0-degree layer strips 252 may be directly stacked on top of each other in the stringer base laminate 902. The width of the stringer base laminate 902 may be at least as wide as the width spanning the flange side edge 906 of the outer wing stringer 272.

[0216] refer to Figure 67 The stringer base laminate 902 has a laminate side edge 904, which gradually decreases in angle with a taper 608 of no more than 30 degrees relative to the inner surface of the corresponding upper skin panel 240 or lower skin panel 242 to which the stringer base laminate 902 is directly or indirectly connected. Preferably, the taper 608 is between approximately 15 and 30 degrees to reduce imprints in the skin panels 240, 242 as described above.

[0217] refer to Figure 65 The outer wing upper skin panel 240 and / or outer wing lower skin panel 242 may each have one or more 0-degree layer strips 252 interwoven in the area layers 302 of the skin panels 240, 242 at one or more mounting locations of the outer wing stringers 272. Each 0-degree layer strip 252 may have a width at least as wide as the stringer base laminate 902 or at least as wide as the width spanning the stringer flange 276, whichever is greater.

[0218] like Figure 65 As shown, the 0-degree layer strip 252 may be biased toward the inner surface of the skin panels 240, 242. However, the 0-degree layer strip 252 may be located within the stack of area layers 302 such that the 0-degree layer strip 252 does not define the inner surface of the skin panels 240, 242. Instead, one of the area layers 302 preferably defines the inner surface of the skin panels 240, 242. Although Figure 65An example of a skin panel 240, 242 with three 0-degree layer strips 252 interwoven within area layer 302 directly below the stringer mounting location is shown; however, skin panels 240, 242 may include any number of 0-degree layer strips 252 at each stringer mounting location. For example, skin panels 240, 242 may include 1-10 0-degree layer strips 252 at each stringer mounting location. As described below, the number of 0-degree layer strips 252 at each stringer mounting location can be determined by the bending stiffness requirements at the spanwise location of the outer wing structure 210. The 0-degree layer strips 252 interwoven within area layer 302 may have a constant width along the spanwise direction.

[0219] refer to Figures 68 to 69 The alternative construction of the outer wing stringer 272 is shown. Figure 68 An example is shown of an outer wing stringer 272 having a stringer base laminate 902 consisting only of a 0-degree layer strip 252, which is connected to a lower skin panel 242 that does not have a 0-degree layer strip 252. Figure 69 An example of the outer wing stringer 272 is shown, wherein the stringer base laminate 902 and the lower skin panel 242 do not have 0-degree layer strips 252. In each example, the stringer cover 910 includes interlaced 0-degree layer strips 252 to increase the bending stiffness of the outer wing stringer 272.

[0220] refer to Figure 70 The diagram shows a spanwise cross-sectional view of an example of an outer wing structure 210, illustrating a ply reduction 264 in the 0-degree layer strip 252 of the stringer canopy 910. In some instances, the stringer layers 908 comprising the outer wing stringers 272 and the stringer base laminate 902 can be continuous and constant along the spanwise direction from the outer wing transverse rib 232 toward the wingtip 126. While the outer wing stringers 272 can each have a constant cross-sectional shape and dimensions, the spanwise position of the ply reduction 264 in the 0-degree layer strip 252 of the stringer canopy 910 can be selected to produce a desired bending stiffness distribution of the outer wing stringers 272 along the spanwise direction. The position of the ply reduction 264 can depend on the predicted aerodynamic wing loads on the outer wing structure 210 during flight.

[0221] Figure 70The diagram also illustrates a ply reduction 264 optionally included in the 0-degree layer strips 252 within the stringer base laminate 902. Additionally, ply reduction 264 optionally interwoven within the 0-degree layer strips 252 of the outer wing lower skin panel 242 or the outer wing upper skin panel 240 is shown. As described above, the area layers 302 constituting each skin panel 240, 242 are continuous and constant along the spanwise direction, resulting in a constant thickness for the skin panels 240, 242, which simplifies manufacturing. Advantageously, the 0-degree layer strips 252 optionally included in the skin panels 240, 242 may have one or more ply reductions 264 to facilitate a gradual decrease in the bending stiffness of the outer wing structure 210 along the spanwise direction.

[0222] about Figures 28 to 43 and Figures 48 to 56 The structure of the outer wing 210 shown is as follows: Figures 59 to 70 The illustrated outer wing structure 210 can be configured such that the combination of the upper skin panel 240 and the associated outer wing stringers 272 has a higher axial stiffness in the spanwise direction than the combination of the lower skin panel 242 and the associated outer wing stringers 272. The higher axial stiffness of the upper skin panel 240 (relative to the lower skin panel 242) can accommodate the compressive loads on the upper skin panel 240 during upward bending of the outer wing structure 210.

[0223] Figure 71 Is manufacturing with such Figures 59 to 70 A flowchart of a method 1000 for a wing assembly 120 of an outer wing structure 210 is shown. Method 1000 includes step 1002 of bonding one or more outer wing stringers 272 to an upper outer wing skin panel 240 and / or a lower outer wing skin panel 242 of the outer wing structure 210. As described above, each outer wing stringer 272 extends in the spanwise direction and is composed of stringer layers 908. Each outer wing stringer 272 has a cap-shaped section 900 and is composed of a pair of stringer flanges 276, a pair of stringer webs 278, and a stringer cap 910. The stringer cap 910 includes one or more 0-degree layer strips 252 interwoven within the stringer cap 910.

[0224] Manufacturing with Figures 59 to 70 The method 1000 for the wing assembly 120 of the outer wing structure 210 shown above may include method 500. Figure 46 ) or method 800 ( Figure 58 Any one or more of the above steps, except that the 0-degree strip group 250 is not interwoven into the skin panels 240, 242 of the outer wing structure 210 (e.g., Figures 41 to 45 Alternatively, the slats 600 can be glued to the skin panels 240, 242 of the outer wing structure 210 (e.g., Figures 54 to 55Method 1000 includes bonding the outer wing stringers 272 to the skin panels 240, 242 of the outer wing structure 210, such as... Figures 59 to 70 As shown.

[0225] Before bonding the outer wing stringers 272 to the skin panels 240, 242, method 1000 may include laying multiple fabric layers 304 to produce the outer wing stringers 272. Each outer wing stringer 272 has a cap-shaped cross-section. Each outer wing stringer 272 may be laid with only fabric layers 304 and may lack unidirectional layers 306, except for the 0-degree layer strips 252 in the stringer cover 910 and / or the stringer base laminate 902. The stringer layers 908 (except for the 0-degree layer strips 252) may be constant from the outer wing transverse ribs 232 toward the wingtip 126 (i.e., no ply reduction 264). The step of laying the outer wing stringers 272 may include interlacing one or more 0-degree layer strips 252 within the stringer layers 908 of the stringer cover 910. Although Figure 65 An example of a stringer cap 910 with four 0-degree layer strips 252 is shown, but the stringer cap 910 may include any number of 0-degree layer strips 252. For example, the stringer cap 910 may include any number of 1-10 0-degree layer strips 252. The number of 0-degree layer strips 252 can be determined by the bending stiffness requirements of the skin panels 240, 242. Figures 65 to 66 As shown and as described above, each of the 0-degree layer strips 252 can be confined to the stringer cover 910 and can be excluded from the radial portion 182 and the stringer web 278 on opposite sides of the stringer cover 910. The step of weaving the 0-degree layer strips 252 within the stringer cover 910 may include incorporating one or more ply reductions 264 in the 0-degree layer strips 252 of the stringer cover 910 to reduce the bending stiffness of the outer wing structure 210 along the spanwise direction. As described below, the location of the ply reductions 264 can be determined by the bending stiffness requirements at different locations along the spanwise direction of the outer wing structure 210.

[0226] Preferably, the outer wing stringers 272 on each of the outer wing upper skin panel 240 and / or outer wing lower skin panel 242 have the same construction. For example, the outer wing stringers 272 on the outer wing lower skin panel 242 may all have the same cross-sectional dimensions, the same cross-sectional shape, the same layer stacking order, and / or the same composite material. In this respect, except for the length of the outer wing stringers 272, the construction of the outer wing stringers 272 on each skin panel 240, 242 can be identical, which simplifies manufacturing and also keeps the chordal mass distribution of the outer wing stringers 272 within 10% of each other. The outer wing stringers 272 on the outer wing upper skin panel 240 may have the same arrangement as the outer wing stringers 272 on the outer wing lower skin panel 242.

[0227] In some instances, step 1002 of bonding the outer wing stringers 272 to the upper and lower wing skin panels 240 and 242 may include bonding a stringer base laminate 902 to the upper and / or lower wing skin panels 240, such that the stringer flange 276 of each outer wing stringer 272 is mounted on the stringer base laminate 902. Figure 65 , Figure 68 and Figure 69 As shown, the stringer base laminate 902 can be positioned between the stringer flange 276 and the skin panels (i.e., the outer wing upper skin panel 240 and / or the outer wing lower skin panel 242). The stringer base laminate 902 may include one or more 0-degree layer strips 252 interwoven within the stringer layers 908 of the stringer base laminate 902. The width of each stringer base laminate 902 is at least as wide as the width spanning the flange side edge 906. To prevent resin cracking, the stringer base laminate 902 may include no more than four 0-degree layer strips 252 directly stacked on top of each other.

[0228] The stringer base laminate 902 may include a unidirectional layer 306 and may not have a fabric layer 304. As described above, the unidirectional layer 306 of the stringer base laminate 902 may include a 0-degree layer and a non-0-degree layer. The non-0-degree layer may include a 90-degree layer, a +45-degree layer, and a -45-degree layer, or a unidirectional layer 306 with other fiber orientations. As described above, the 0-degree layer has fibers oriented parallel to the longitudinal direction of the outer wing stringer 272. The 90-degree layer has fibers oriented perpendicular to the longitudinal direction of the outer wing stringer 272. The +45-degree layer has fibers oriented at a 45-degree angle relative to the longitudinal direction of the outer wing stringer 272.

[0229] Method 1000 may include forming each of the opposing laminate side edges 904 of the stringer base laminate 902 with an edge taper angle 608 of no more than 30 degrees relative to the inner surface of the respective upper skin panel 240 or lower skin panel 242 to which the stringer base laminate 902 is bonded. Preferably, the edge taper angle 608 is between approximately 15 and 30 degrees. The shallow edge taper angle 608 in the base laminate can reduce or prevent imprinting on the inner surface of the skin panels 240, 242 during the bonding of the stringer base laminate 902 to the skin panels 240, 242. Furthermore, in this respect, method 1000 may include forming the flange-side edge 906 of each outer wing stringer 272 with an edge taper angle 608 of no more than 30 degrees (e.g., preferably between 15 and 30 degrees) to prevent imprints in the stringer base laminate 902, or, for example, to prevent imprints in the skin panels 240, 242 to which the stringer flange 276 may be glued, in the case that the stringer base laminate 902 is omitted.

[0230] Step 1002, which involves bonding the outer wing stringer 272 to the upper wing skin panel 240 and / or the lower wing skin panel 242, may include secondary bonding of the stringer flange 276 and / or the stringer base laminate 902 (e.g., in a cured state) of the outer wing stringer 272 to the cured skin panels 240, 242 via an adhesive layer (not shown). Alternatively, step 1002 may include co-bonding the outer wing stringer 272 (e.g., in a cured or uncured state) to the skin panels 240, 242, which may be in a cured or uncured state. In yet another example, step 1002 may include co-curing the outer wing stringer 272 (e.g., in a cured state) and the skin panels 240, 242 (e.g., in an uncured state).

[0231] Step 1002, which involves bonding the outer wing stringers 272 to the upper and / or lower skin panels of the outer wing, may further include positioning the outer wing stringers 272 on each skin panel 240, 242 at a chordal spacing 256 of 6-11 inches between stringer centerlines 912. More preferably, the chordal spacing 256 is 8-10 inches. Preferably, the chordal spacing 256 between the outer wing stringers 272 is uniform in order to maintain a substantially uniform chordal mass distribution of the outer wing structure 210. In some instances, the outer wing stringers 272 may be positioned on the skin panels 240, 242 in a generally parallel arrangement (e.g., ±30 degrees) to each other. The outer wing stringers 272 in each outer wing structure 210 may be aligned with the center wing stringer 270 of the center wing structure 150, respectively. However, the outer wing stringers 272 may not be aligned with the center wing stringer 270.

[0232] As described above, method 1000 may include laying multiple area layers 302 to produce an upper wing skin panel 240 and a lower wing skin panel 242. As described above, the area layers 302 may be substantially constant along the chord direction between the outer wing's forward sparsity 212 and a rearward sparsity 214, and substantially constant along the spanwise direction between the outer wing's transverse ribs 232 and wingtips 126. To increase the bending stiffness of the outer wing skin panels, method 1000 may further include interlacing one or more 0-degree layer strips 252 within the area layers 302 at one or more stringer mounting locations (i.e., chord positions) of the outer wing stringers 272 on the upper wing skin panel 240 and the lower wing skin panel 242. To avoid or reduce the risk of layer splitting failure modes in the skin panels 240, 242, the 0-degree layer strips 252 within the skin panels 240, 242 may be positioned between area layers 302 that are non-0-degree layers. To prevent resin cracking, no more than four 0-degree layer strips 252 may be stacked close to each other in the skin panels 240, 242.

[0233] To reduce the bending stiffness of the outer wing structure 210 along the spanwise direction, method 1000 may further include incorporating one or more ply reductions 264 in the 0-degree layer strips 252 interwoven within the area layers 302 of the outer wing upper skin panel 240 and / or the outer wing lower skin panel 242. As described above, the spanwise position of the ply reductions 264 in the 0-degree layer strips 252 of (1) the stringer cover 910, (2) the stringer base laminate 902, and / or (3) the skin panels 240, 242 can be based on the desired bending stiffness distribution of the skin panels 240, 242 along the spanwise direction. This stiffness distribution can be based on predicted aerodynamic wing loads on the outer wing structure 210 during flight.

[0234] The step of laying the outer wing stringers 272 may include laying each outer wing stringer 272 to have a stringer bending stiffness of approximately 40-60% (e.g., more preferably 45-55%) of the combined slat-skin portion bending stiffness of the slats 600 and the skin panel portion 614 of the outer wing upper skin panel 240 or the outer wing lower skin panel 242. Figure 62 As shown and described above, the skin panel portion 614 has a panel portion width 618 that extends between midpoints 622 located on opposite sides of the outer wing stringers 272. Each midpoint 622 is located midway between an outer wing stringer 272 and an adjacent outer wing stringer 272. Alternatively, the panel portion width 618 of the skin panel portion 614 may extend between a spar on one side of the outer wing stringer 272 and a midpoint 622 between an outer wing stringer 272 on the opposite side of the outer wing stringer 272 and an adjacent outer wing stringer 272.

[0235] Method 1000 may further include laying the upper outer wing skin panel 240 and associated outer wing stringers 272 in a manner having a combined axial stiffness in the spanwise direction, the combined axial stiffness being higher than that of the lower outer wing skin panel 242 and associated outer wing stringers 272. As described above, the higher axial stiffness of the upper skin panel 240 can be determined by the relatively high compressive load induced in the upper skin panel 240 during the upward bending of the outer wing structure 210, compared to the lower axial stiffness requirement of the lower skin panel 242 due to the increased tensile load bearing capacity of the composite structure caused by the high tensile strength of the fiber relative to the resin compressive strength.

[0236] For an example of an outer wing structure 210 having a fuel tank 128, method 1000 may include attaching the outer wing stringers 272 to the upper outer wing skin panel 240 and / or to the lower outer wing skin panel 242 in such a manner that each outer wing structure 210 has no outer wing stringers 272 beyond the outer end 134 of the fuel tank 128. Furthermore, in this respect, method 1000 may include laying the outer wing stringers 272 such that the stringer cap 910, the stringer base laminate 902, and / or the 0-degree layer strips 252 in the skin panels 240, 242 do not extend beyond the outer end 134 of the fuel tank.

[0237] Step 1004 of method 1000 is substantially similar to step 702 of method 700 described above. For example, step 1004 includes connecting the outer wing upper skin panel 240 and outer wing lower skin panel 242 (and associated outer wing stringers 272) to the outer wing front spars 212, outer wing rear spars 214, and outer wing ribs 234 (including outer wing transverse ribs 232) of the outer wing structure 210 (e.g., Figures 59 to 70 As described above, the upper skin panel 240 and the lower skin panel 242 of the outer wing can be adhesively and / or mechanically fastened to the front wing spars 212 and the rear wing spars 214 of the outer wing, as well as the wing ribs 234.

[0238] Step 1006 of method 1000 is substantially similar to step 706 of method 700 described above. Step 1006 includes attaching an outer wing structure 210 (e.g., Figures 59 to 70 Each of the outer wing inner ends 230 is respectively connected to the opposite outer end 168 of the central wing structure 150 to define a pair of wing joints 172. As described above and Figure 4 As shown, the center wing structure 150 has engine mounting positions on each opposite side of the wing centerline 124. At each wing joint 172, the spars ends 156 of the center wing forward spars 152 and center wing aft spars 154 are located no more inward than the engine centerline 106 associated with the engine mounting positions, and no more than 10 percent of the distance between the engine centerline 106 and the wing centerline 124, further outward from the engine centerline 106.

[0239] Additionally, step 1006 may include any one or more of the aforementioned alternative or additional steps related to step 706. For example, step 1006 may include connecting the outer wing transverse rib 232 of the outer wing structure 210 to the center wing transverse rib 170 of the center wing structure 150, such that the outer wing transverse rib 232 and the center wing transverse rib 170 are positioned as follows: Figure 17The back-to-back arrangement is shown and described above. Step 1006 may further include using a front spar splice plate 216 to mechanically fasten the spar terminal 156 of the center wing front spar 152 to the spar terminal 156 of the outer wing front spar 212, and using a rear spar splice plate 218 to mechanically fasten the spar terminal 156 of the center wing rear spar 154 to the spar terminal 156 of the outer wing rear spar 214, as shown. Figure 5 As shown and described above.

[0240] Figure 72 It is loaded with the features described above Figures 59 to 70 A flowchart of a method 1100 for constructing an outer wing structure 210 and a wing assembly 120. The method includes step 1102 of providing a wing assembly 120 having a pair of outer wing structures 210, each outer wing structure being connected to a central wing structure 150. The central wing structure 150 may be substantially similar to... Figures 1 to 27 The arrangement shown above. The construction of each outer wing structure 210 is similar to that of the above. Figures 59 to 70 The arrangement is shown. Method 1100 may include step 1104 of placing the wing assembly 120 under static ground load conditions, and step 1106 of placing the wing assembly 120 under dynamic load conditions. As described above, the static ground load conditions relate to the aircraft 100 being substantially stationary and under static ground load conditions, under which the wing assembly 120 is subjected to gravity due to the structural mass of the wing assembly 120. The dynamic load conditions relate to the motion of the aircraft 100 and the wing assembly 120 being subjected to one of the following: gravity due to the structural mass of the wing assembly 120, inertial forces due to the structural mass of the wing assembly 120, aerodynamic loads on the wing assembly 120, and / or control surface loads on the wing assembly 120.

[0241] Now for reference Figures 73 to 86 An example of an outer wing structure 210 with an upper outer wing panel assembly 1200 and a lower outer wing panel assembly 1202 is shown. Figures 73 to 79 As shown, each panel assembly 1200, 1202 extends along the wingspan and connects to the outer wing's forward spar 212 and rear spar 214. Each panel assembly 1200, 1202 can extend from the outer wing's transverse rib 232 to the wingtip 126. Figure 76As shown, each panel assembly 1200, 1202 includes an outer skin panel 1204 and a cap-shaped section panel 1206. The outer skin panel 1204 consists of an area layer 302 and has a constant panel thickness along the spanwise direction from the outer wing transverse rib 232 to the wingtip 126. The cap-shaped section panel 1206 is attached to the inner side of the outer skin panel 1204. The cap-shaped section panel 1206 includes a plurality of generally parallel (e.g., ±30 degrees) cap-shaped segments 1210 spaced apart from each other in the chord direction. The cap-shaped segments 1210 are integrally formed with the cap-shaped section panel 1206 and extend in the spanwise direction. In addition to the upper panel assembly 1200 and the lower panel assembly 1202, Figures 73 to 79 The outer wing structure 210 is constructed in a manner similar to the above. Figures 28 to 43 210, outer wing structure Figures 48 to 56 210 outer wing structure and Figures 59 to 70 The outer wing structure 210.

[0242] For example, Figures 73 to 79 The outer wing structure 210 includes a front wing sparb 212 and a rear wing sparb 214, which are respectively connected to the front wing sparb 152 of the center wing. Figure 5 ) and center wing rear spars 154 ( Figure 5 ), to limit the connection of the outer wing structure 210 to the central wing structure 150 ( Figure 5 ) Wing joint 172 ( Figure 5 The wing joint 172 may include a front wing spars splice plate 216. Figure 5 ) and rear wing beam splice plate 218 ( Figure 5 ), used to connect the front spar 212 and rear spar 214 of the outer wing structure 210 to the front spar 152 and rear spar 154 of the center wing structure 150, as described above. Additionally, Figures 73 to 79 The outer wing structure 210 includes an outer wing transverse rib 232 for connection to the center wing transverse rib 170 as described above. Figure 5 The upper wing panel assembly 1200 and the lower wing panel assembly 1202 are respectively connected to the front wing spars 212, the rear wing spars 214 and the wing ribs 234, as described below.

[0243] refer to Figures 75 to 77The diagram shows a cross-sectional view of the outer wing structure 210, illustrating examples of the upper wing panel assembly 1200 and the lower wing panel assembly 1202. As described above, the upper wing panel assembly 1200 and the lower wing panel assembly 1202 each include an outer skin panel 1204 and a cap-shaped section panel 1206 as described above. The area layer 302 of the outer skin panel 1204 may include 10-70 unidirectional layers 306 and may not have a fabric layer 304. In some examples, the combination of the outer skin panel 1204 and the cap-shaped section panel 1206 may have a total of 32 composite layers 300, resulting in a panel thickness 194 that can provide a level of structural integrity that meets the flammability requirements of certain aviation regulatory agencies such as the Federal Aviation Administration (FAA). The unidirectional layers 306 of the outer skin panel 1204 may include 0-degree layers and 90-degree layers.

[0244] Still referencing Figures 75 to 77 Each cap-shaped section panel 1206 is attached to the inner side of the outer skin panel 1204. As described above, the cap-shaped section panel 1206 extends between the outer wing's forward spar 212 and the outer wing's rear spar 214. For example, the leading edge of the cap-shaped section panel 1206 may terminate directly in front of the outer flange 160 of the outer spar of the outer wing's forward spar 212. Similarly, the trailing edge of the cap-shaped section panel 1206 may terminate directly behind the outer flange 160 of the outer spar of the outer wing's rear spar 214. The leading and trailing edges of the cap-shaped section panel 1206 may be gradually tapered to minimize stress concentrations that might otherwise occur in the leading and trailing edges of the cap-shaped section panel 1206. In the example shown, the outer skin panels 1204 of the upper panel assembly 1200 and the lower panel assembly 1202 may extend forward of the outer wing's forward spar 212 and behind the outer wing's rear spar 214.

[0245] The cap-shaped section panel 1206 is composed of a cap-shaped section panel layer 1208. The cap-shaped section panel layer 1208 may include a fabric layer 304 and may not include the unidirectional layer 306, except for the 0-degree layer strip 252 in the cap-shaped section cover 1216. The cap-shaped segments 1210 of the cap-shaped section panel 1206 are spaced apart from each other and are interconnected by cap-shaped section flanges 1212, resulting in the cap-shaped section panel 1206 having a corrugated cross-sectional shape. Each cap-shaped segment 1210 of the cap-shaped section panel 1206 may have the same cross-sectional dimensions and the same cross-sectional shape. The cap-shaped section panel 1206 can increase the bending stiffness of the outer wing structure 210. Each panel assembly 1200, 1202 may include a foam member 1228 sandwiched between the outer skin panel 1204 and the cap-shaped section panel 1206 at the location of each cap-shaped segment 1210, as described below. Each panel assembly 1200, 1202 may also include a slat 600 consisting of a 0-degree slat layer 602 sandwiched between the outer skin panel 1204 and each cap-shaped section flange 1212, as also described below.

[0246] Although the cap-shaped cross-section panel 1206 of each panel assembly 1200, 1202 is shown as (e.g., Figure 75 The outer wing structure 1206 has four cap-shaped segments 1210, but the cap-shaped section panel 1206 may include any number of cap-shaped segments 1210. For example, the cap-shaped section panel 1206 may include only a single cap-shaped segment 1210, or the cap-shaped section panel 1206 may include two or more cap-shaped segments 1210. Each cap-shaped segment 1210 has a cap-shaped section centerline 1220 extending along the longitudinal direction of the cap-shaped segment 1210. The cap-shaped segments 1210 may be spaced apart from each other by a chordal spacing 256 of 6-11 inches, and more preferably by a chordal spacing 256 of 8-10 inches. The chordal spacing 256 may be uniform among the cap-shaped segments 1210. The cap-shaped segments 1210 in each outer wing structure 210 may be respectively associated with the center wing structure 150. Figure 5 ) center wing stringer 270 ( Figure 5 Alignment. Alternatively, the cap-shaped section 1210 in the outer wing structure 210 may not be aligned with the center wing stringer 270.

[0247] like Figure 75 As illustrated in the example, the cap-shaped sections 1210 of the cap-shaped section panel 1206 can all have the same construction, such as the same cross-sectional shape and / or the same cross-sectional dimensions, which simplifies manufacturing. Furthermore, the common cross-sectional shape and / or dimensions of the cap-shaped sections 1210 can keep the chordal mass distribution of the outer wing structure 210 within 10% at all chordal locations. Additionally, the cross-sectional shape and / or cross-sectional dimensions of each cap-shaped section 1210 can be constant along the longitudinal direction.

[0248] Brief Reference Figures 78 to 79 The diagram shows an outer wing rib 234 extending between the front wing spars 212 and the rear wing spars 214 of the outer wing. The outer wing rib 234 has rib cutouts 268 for receiving cap-shaped sections 1210 of the upper wing panel assembly 1200 and the lower wing panel assembly 1202. The outer flanges 178 of the outer wing rib 234 can be bonded and / or mechanically fastened to the upper wing panel assembly 1200 and the lower wing panel assembly 1202. Additionally, the outer wing rib 234 can be bonded and / or mechanically fastened to the spars outer flanges 160 of the front wing spars 212 and the rear wing spars 214 of the outer wing.

[0249] refer to Figure 80 Each cap-shaped segment 1210 may be symmetrical about the vertical axis 1222 of the cap-shaped section. Each cap-shaped segment 1210 may have a cap-shaped section width 1224 between 3 and 6 inches, more preferably between 4 and 5 inches. The cap-shaped section width 1224 may be measured between the intersection of the cap-shaped section flange 1212 and the cap-shaped section web 1214 on each side of the axis 1222 of the cap-shaped section. In addition, each cap-shaped segment 1210 may have a cap-shaped section height 1226 not greater than 3 inches.

[0250] refer to Figure 80 and Figure 82 The cap-shaped section panel 1206 includes a plurality of cap-shaped section panel layers 1208 and is formed in a corrugated shape defined by the plurality of cap-shaped segments 1210. Each cap-shaped segment 1210 includes a pair of cap-shaped section flanges 1212. Adjacent pairs of cap-shaped segments 1210 share a common cap-shaped section flange 1212. In addition, each cap-shaped segment 1210 includes a pair of cap-shaped section webs 1214. Each cap-shaped section web 1214 is connected to the cap-shaped section flange 1212 via a radius portion 182. Each cap-shaped segment 1210 also includes a cap-shaped section cap 1216 interconnected with the cap-shaped section webs 1214. Each cap-shaped section web 1214 is interconnected with the cap-shaped section cap via a radius portion 182.

[0251] Each cap 1216 comprises one or more 0-degree layer strips 252 interwoven within the cap 1208 of the cap 1206. Each 0-degree layer strip 252 has fibers aligned with the wingspan or longitudinal direction of the cap segment 1210. As described above, the 0-degree layer strips 252 may be located at approximately the mid-plane of the cap 1216. Each 0-degree layer strip 252 in the cap 1216 may extend across the entire width of the cap 1216. To avoid warping due to thermal cycling during curing, each 0-degree layer strip 252 is confined to the cap 1216 and excluded from the radial portion 182 on each side of the cap 1216. To prevent resin cracking, no more than four 0-degree layer strips 252 are stacked directly on top of each other in the cap 1216. The interwoven 0-degree layer strips 252 increase the bending stiffness of panel assemblies 1200 and 1202.

[0252] refer to Figure 83 Each 0-degree layer strip 252 in the cap-shaped section cover 1216 may have one or more ply reductions 264 to gradually decrease the bending stiffness of the outer wing structure 210 along the spanwise direction. However, the cap-shaped section panel layer 1208 (excluding the 0-degree layer strip 252) may be constant or continuous in the spanwise direction. The spanwise position of the ply reductions 264 of the 0-degree layer strips 252 in the cap-shaped section cover 1216 can be selected to produce a desired bending stiffness distribution of the outer wing structure 210 along the spanwise direction, and can be based on predicted wing loads on the outer wing structure 210 during flight.

[0253] refer to Figures 76 to 77 and Figures 80 to 81 Each panel assembly 1200, 1202 may include multiple slats 600 extending in the spanwise or longitudinal direction of the cap-shaped section 1210. The cap-shaped slat layer 602 constituting the slats 600 may be continuous and constant along the spanwise direction from the outer wing transverse rib 232 toward the wingtip 126. Although Figure 80 An example of a slat 600 having three slat layers 602 is shown, but a slat 600 may contain any number of slat layers 602. Slat layers 602 may include 0-degree layers and / or 0-degree layers. If non-0-degree layers are included, 0-degree layers may comprise the maximum percentage of slat layers 602 in each slat 600.

[0254] Each panel assembly 1200, 1202 may include a slat 600 on opposite sides of each cap-shaped section 1210. Each slat 600 may be embedded between the cap-shaped flange 1212 of the outer skin panel 1204 and the cap-shaped section panel 1206. Each slat 600 may be laid and may be secondary bonded (e.g., via an adhesive layer), co-cured, or co-bonded to the outer skin panel 1204 and the cap-shaped section panel 1206. Figure 84 As shown, the slats 600 of the hat-shaped cross-section panel 1206 may include ply reductions 264 in one or more slat layers 602. Each ply reduction 264 may be located at a predetermined spanwise position corresponding to the desired stiffness requirement at a spanwise position of the outer wing structure 210.

[0255] refer to Figure 81 Each slat 600 has opposing slat side edges 606. At least some of the slat side edges 606 of the slats 600 may taper relative to the inner surface of the corresponding outer skin panel 1204 at an edge taper angle 608 not greater than 30 degrees, resulting in a generally trapezoidal cross-sectional shape for each slat 600. More preferably, the edge taper angle 608 is between approximately 15 and 30 degrees. By forming the slat side edges 606 with an edge taper angle 608, imprints on the inner surface of the outer skin panel 1204 can be reduced or prevented.

[0256] refer to Figures 80 to 81 The inner surface of the outer skin panel 1204 and the slat side edges 606 of adjacent paired slats 600 each define a recess 1230 on the opposite side of each slat 600. The recess 1230 is a portion of a cap-shaped section cavity 1218 defined by the cap-shaped section web 1214 and cap-shaped section cap 1216 of each cap-shaped segment 1210. Each cap-shaped section cavity 1218 (i.e., the recess 1230) can be filled with a foam member 1228. Each foam member 1228 may have a cross-sectional shape complementary to the cross-sectional shape of the cap-shaped section cavity 1218, such that each foam member 1228 occupies the entire cap-shaped section cavity 1218. By filling the entire cap-shaped section cavity 1218, the foam member 1228 can meet the flammability requirements of certain aviation regulatory agencies (e.g., the FAA). Additionally, the foam member 1228 can increase the overall flexural stiffness of the panel assemblies 1200, 1202. Furthermore, foam component 1228 can be used as a tool for laying and / or curing cap-shaped sections 1210 on the cap-shaped section panel 1206 during its installation. Foam component 1228 can be made of closed-cell foam material (e.g., Rohacell). TM )composition.

[0257] refer to Figure 83 For an example of an outer wing structure 210 including a fuel tank 128 (e.g., Figures 73 to 74The upper panel assembly 1200 and / or the lower panel assembly 1202 can be configured such that the cap-shaped section panel 1206 terminates at the outer end 134 of the fuel tank 128 due to reduced stiffness requirements on the outer side of the outer end 134. The reduced stiffness requirement may be due to a reduction in inertial load of the outer wing structure 210 on the outer side of the fuel tank 128. However, the outer skin panel 1204 can extend continuously from the outer wing transverse rib 232 to the wingtip 126. For panel assemblies 1200 and 1202 including foam members 1228 and / or slats 600, the slats 600 and foam members 1228 can terminate at the same location as the end of the cap-shaped section panel 1206 (e.g., at the outer end 134).

[0258] Brief Reference Figure 76 The diagram shows a cross-section of a portion of the outer wing structure 210, illustrating the short cap-shaped section height 1226 of each cap-shaped segment 1210 relative to the large height of a conventional stringer (not shown). Due to the short cap-shaped section height 1226, each panel assembly 1200, 1202 has its area centroid closer to the area centroid 260 of the skin panels 240, 242 than that of skin panels with conventional cap-shaped stringers. The distance between the area centroids 260 of the upper panel assembly 1200 and the lower panel assembly 1202 corresponds to... Figures 73 to 80 The outer wing structure 210 has a larger effective wing thickness 262 compared to the smaller effective wing thickness of wing assemblies using conventional stringers. Figures 73 to 80 The increased effective wing thickness 262 of the outer wing structure 210 leads to an increase in the bending load bearing efficiency of the outer wing structure 210, and this allows for a reduction in the structural mass of the outer wing structure 210.

[0259] The upper panel assembly 1200 can have a higher axial stiffness in the spanwise direction than the lower panel assembly 1202. This increased axial stiffness requirement for the upper panel assembly 1200 can accommodate compressive loads in the upper panel assembly 1200 during upward bending of the outer wing structure 210. To this end, the upper panel assembly 1200 can have an increased number of 0-degree layer strips in the cap-shaped section cover 1216 and / or in the slats 600, and / or an increased number of cap-shaped section panel layers 1208 or larger cross-sectional dimensions comprising cap-shaped sections 1210 of the cap-shaped section panel 1206 of the upper panel assembly 1200.

[0260] Figure 85 It is to manufacture products with the features described above. Figures 73 to 84A flowchart of a method 1300 for a composite wing assembly 120 of the outer wing structure 210 is shown. Method 1300 includes step 1302 of bonding a cap-shaped section panel 1206 to the outer skin panels 1204 of the outer wing upper panel assembly 1200 and the outer wing lower panel assembly 1202 of the outer wing structure 210. As described above, each outer skin panel 1204 of the outer wing upper panel assembly and the outer wing lower panel assembly consists of an area layer 302 having a panel thickness that is constant along the wingspan direction. The cap-shaped section panel 1206 may also have a panel thickness that is constant along the wingspan direction. Additionally, the cap-shaped section panel 1206 has a plurality of generally parallel cap-shaped segments 1210, each cap-shaped section extending in the wingspan direction.

[0261] Manufacturing with Figures 73 to 84 The method 1300 for the wing assembly 120 of the outer wing structure 210 may include method 500. Figure 46 Method 800 Figure 58 ) or any one or more of the above steps of method 1000, except that the 0-degree strip group 250 is not interwoven into the skin panels 240, 242 of the outer wing structure 210 (e.g., Figures 41 to 45 Or use 600 slats (for example, Figures 54 to 55 ) or outer wing spars 272 (e.g., Figure 62 and Figures 68 to 69 Method 1300 includes bonding a cap-shaped section panel 1206 to an outer wing panel 1204 to produce an upper outer wing panel assembly 1200 or a lower outer wing panel assembly 1202, such as... Figures 76 to 77 As shown.

[0262] Before bonding the cap-shaped section panel 1206 to the outer skin panel 1204, method 1300 may include laying a certain amount of area layer 302 to produce the outer skin panel 1204. Area layer 302 may include a unidirectional layer 306 and may not include the fabric layer 304. The step of laying area layer 302 may include laying 0-degree and 90-degree layers at a ratio of 0-degree to 90-degree layers between 1.75 and 2.5 to provide higher stiffness in the spanwise direction relative to stiffness in the chord direction. However, the outer skin panel 1204 may be laid at any ratio of 0-degree to 90-degree layers.

[0263] Method 1300 may further include laying multiple cap-shaped section panel layers 1208 to produce a cap-shaped section panel 1206. Method 1300 may include using a fabric layer 304 to lay the cap-shaped section panel 1206 to form a corrugated shape defined by spaced-apart cap-shaped segments 1210. The cap-shaped section panel 1206 may be without the unidirectional layer 306 to avoid warping and / or wrinkling that may occur during the curing of the non-planar composite laminate laid by the unidirectional layer 306. As described above, the cap-shaped segments 1210 of the cap-shaped section panel 1206 may all have the same construction, which can keep the chordal mass distribution of the outer wing structure 210 within 10% at all chordal locations.

[0264] The step of laying the cap-shaped section panel layer 1208 to produce the cap-shaped section panel 1206 may include spacing the cap-shaped segments 1210 apart from each other with a chordal spacing 256 of 6-11 inches between the cap-shaped section centerlines 1220. Each cap-shaped segment 1210 may be laid symmetrically about the vertical axis 1222 of the cap-shaped section. Additionally, as described above, the cap-shaped segments 1210 in each outer wing structure 210 may be aligned with the center wing stringer 270 of the center wing structure 150. The step of laying the cap-shaped section panel 1206 may also include laying each cap-shaped segment 1210 wherein the cap-shaped section width 1224 is between 3-6 inches, and / or the cap-shaped section height 1226 is no greater than 3 inches, to keep the area centroid close to the outer skin panel 1204.

[0265] The step of laying multiple cap-section panel layers 1208 to produce cap-section panels 1206 may include interlacing one or more 0-degree layer strips 252 within the cap-section panel layer 1208 at each of the corresponding multiple cap-section caps 1216 of the multiple cap-section segments 1210. The 0-degree layer strips 252 may interlace at approximately the mid-plane of each cap-section cap 1216. As described above, the interlaced 0-degree layer strips 252 may contribute to providing flexural stiffness by the cap-section segments 1210.

[0266] The step of laying the cap-shaped section 1210 may also include incorporating one or more ply reductions 264 in the 0-degree layer strip 252 of the cap-shaped section cover 1216 as a way to gradually reduce the bending stiffness of the outer wing structure 210 along the span direction. Although the 0-degree layer strip 252 includes one or more ply reductions 264, the cap-shaped section panel layer 1208 defining the cap-shaped section panel 1206 can be constant and continuous in the span direction from the outer wing transverse rib 232 toward the wingtip 126.

[0267] In some instances, method 1300 may further include laying a plurality of slat layers 602 on the inner surface of the outer skin panel 1204 to produce a plurality of slats 600, each slat extending in the spanwise direction and respectively located between a plurality of cap-shaped segments 1210. The step of laying the slat layers 602 may include laying a plurality of 0-degree layers and non-0-degree layers to produce the slats 600. As described above, the 0-degree layers may comprise the largest percentage of the slat layers 602 in the slats 600. The slat layers 602 may be planar in shape and therefore may be laid by unidirectional layers 306 comprising 0-degree layers and non-0-degree layers. After laying the slats 600, method 1300 may include positioning the slats 600 between the inner surface of the cap-shaped flange 1212 and the inner surface of the outer skin panel 1204, and bonding the cap-shaped panel 1206 to the slats 600 and the outer skin panel 1204, such that the slats 600 are trapped between the outer skin panel 1204 and the cap-shaped panel 1206. Advantageously, the slats 600 can increase the bending stiffness of the upper panel assembly 1200 and the lower panel assembly 1202.

[0268] The step of laying the slat layer 602 may include laying each slat 600 having opposing slat side edges 606 that taper relative to the inner surface of the corresponding outer skin panel 1204 at an edge taper angle 608 not greater than 30 degrees, and this may result in... Figure 81 The cross-sectional shape of each slat 600 is shown as approximately trapezoidal. The steps of laying the slat layer 602 to form the slats 600 and positioning the slats 600 on the outer skin panel 1204 can result in recesses 1230 defined between adjacent pairs of slats 600, as shown. Figure 81 As shown. As described above, each recess 1230 may form part of a cap-shaped cross-sectional cavity 1218 that may optionally be filled with a foam member 1228.

[0269] Method 1300 may further include incorporating one or more ply reduction steps 264 in the slat layer 602 as a way of gradually reducing the bending stiffness of the outer wing structure 210 along the span direction. As described above, each of the slats 600 may have a slat width 604 between 3 and 6 inches. Figures 76 to 77 As shown, the slat width 604 can be determined by the hat-shaped section width 1224 of each hat-shaped segment 1210 and by the chordal spacing 256 of the hat-shaped segments 1210.

[0270] As described above, the outer wing upper panel assembly 1200 and / or outer wing lower panel assembly 1202 may include foam members 1228. In this regard, method 1300 may include laying a cap-shaped section panel 1206 onto a plurality of foam members 1228 to create a plurality of cap-shaped segments 1210. As described above, the profile of the foam members 1228 may be complementary to the desired configuration of the cap-shaped segments 1210 of the cap-shaped section panel 1206. Each foam member 1228 may serve as a lay-up mandrel and / or a curing mandrel on which the cap-shaped section panel layer 1208 may be laid. After the cap-shaped section panel 1206 has cured, the foam members 1228 may be retained within the cap-shaped section cavity 1218, and the method may further include bonding the cap-shaped section panel 1206 to the outer skin panel 1204 such that the foam members 1228 are trapped therebetween. Alternatively, the foam component 1228 can be removed from each cap-shaped cavity 1218 before the cap-shaped section panel 1206 is bonded to the outer skin panel 1204.

[0271] For example, when the outer wing structure 210 is configured to include a fuel tank 128, method 1300 may include laying each cap-shaped section panel 1206 such that the cap-shaped section panel 1206 does not extend further outward than the outer end 134 of the tank. In this respect, each outer wing structure 210 may not have a cap-shaped section panel 1206 beyond the outer end 134 of the tank (e.g., including slats 600 and / or foam members 1228).

[0272] Step 1304 of method 1300 may be substantially similar to step 704 of method 700 and / or step 1004 of method 1000 described above. For example, step 1304 includes connecting the outer wing upper panel assembly 1200 and the outer wing lower panel assembly 1202 to the outer wing front spar 212 and outer wing rear spar 214 of the outer wing structure 210 (e.g., Figures 73 to 74 The upper wing panel assembly 1200 and the lower wing panel assembly 1202 can be bonded and / or mechanically fastened to the front wing spars 212, the rear wing spars 214, and the wing ribs 234, including the wing transverse ribs 232.

[0273] Step 1306 of method 1300 may be substantially similar to step 706 of method 700 and / or step 1006 of method 1000 described above. For example, step 1306 includes attaching an outer wing structure 210 (e.g., Figures 73 to 74 Each of the outer wing inner ends 230 is respectively connected to the opposite outer end 168 of the central wing structure 150 to define a pair of wing joints 172. As described above and Figure 4As shown, the center wing structure 150 has engine mounting positions on each opposite side of the wing centerline 124. The sparsity ends 156 of the center wing forward spars 152 and center wing aft spars 154 at each wing joint 172 are located no more inward than the engine centerline 106 associated with the engine mounting positions, and no more than 10 percent of the distance between the engine centerline 106 and the wing centerline 124, further outward from the engine centerline 106.

[0274] Figure 86 It is loaded with the features described above Figures 73 to 84 A flowchart of a method 1400 for constructing a composite wing assembly 120 (e.g., applying a load) of an outer wing structure 210, as shown. Method 1400 includes step 1402 of providing a wing assembly 120 having a pair of outer wing structures 210, each outer wing structure being connected to a central wing structure 150. Each outer wing structure 210 can be constructed similar to... Figures 73 to 74 The layout shown.

[0275] Method 1400 includes steps 1404 of placing the wing assembly 120 under static ground load conditions and 1406 of placing the wing assembly 120 under dynamic load conditions. As described above, the static ground load conditions relate to the aircraft 100 being substantially stationary and under static ground load, under which the wing assembly 120 is subjected to gravity due to its structural mass. The dynamic load conditions relate to the motion of the aircraft 100 and the wing assembly 120 being subjected to gravity due to its structural mass, inertial forces due to its structural mass, aerodynamic loads on the wing assembly 120, and / or control surface loads on the wing assembly 120.

[0276] Clause A1: A wing assembly for an aircraft, comprising a center wing structure including a pair of outer wing ends and a pair of engine mounting positions located on opposite sides of the wing centerline; a center wing forward sparb and a center wing aft sparb, each having a sparb termination at each of the outer wing ends; a pair of outer wing structures, each outer wing structure comprising: an outer wing forward sparb and an outer wing aft sparb configured to be connected to the center wing forward sparb and the center wing aft sparb, respectively, to define wing joints connecting the outer wing structures to the center wing structure; and wherein the center wing structure is configured such that the sparb terminations of the center wing forward sparb and the center wing aft sparb are located at each wing joint no more inward than the engine centerline associated with the engine mounting position, and no more outward than 10 percent of the distance between the engine centerline and the wing centerline.

[0277] Clause A2: The wing assembly according to Clause A1 further includes an upper wing skin panel and a lower wing skin panel connected to the outer wing's front wing spars and rear wing spars; wherein at least one of the upper wing skin panel and the lower wing skin panel comprises an area layer having a constant panel thickness along the wingspan direction from the inner end of the outer wing toward the wingtip; and at least one of the upper wing skin panel and the lower wing skin panel has one or more 0-degree stripe groups, each stripe group extending in the wingspan direction, and each stripe group comprising one or more 0-degree layer stripes interwoven within the area layer.

[0278] Clause A3: A wing assembly as described in Clause A1 or A2, wherein each wing joint includes a front spar splice plate connecting the center wing front spar to the outer wing front spar; and a rear spar splice plate connecting the center wing rear spar to the outer wing rear spar.

[0279] Clause A4: A wing assembly according to any one of Clauses A1 to A3, wherein the center wing structure includes a center wing transverse rib located at the outer end of each center wing; each outer wing structure includes an outer wing transverse rib located at the inner end of the outer wing; and at each wing joint, the outer wing transverse rib and the center wing transverse rib are configured to be arranged back-to-back and mechanically fastened to each other.

[0280] Clause A5: The wing assembly according to any one of Clauses A1 to A4 further includes a center wing upper skin panel and a center wing lower skin panel, which are connected to the center wing front spars and the center wing rear spars; an outer wing upper skin panel and an outer wing lower skin panel, which are connected to the outer wing front spars and the outer wing rear spars; wherein at least one of the outer wing structure and the center wing structure has one or more wing stringers, each wing stringer being connected to one of the upper skin panels and the lower skin panels of at least one of the outer wing structure and the center wing structure. Each wing stringer has a stringer flange, at least one stringer web extending outward from the stringer flange, and a stringer transition at the stringer end, the height of the stringer web gradually decreasing along the stringer transition; and at least one of the upper and lower skin panels of the outer wing structure and the center wing structure has a partially raised area of ​​composite layer on the inner side of the skin panel at the chordal position of the stringer flange of at least one wing stringer in the outer wing structure and the center wing structure, the partially raised area being located near one or more stringer ends.

[0281] Clause A6: The wing assembly according to Clause A5, wherein the panel thickness of at least one of the upper skin panel and the lower skin panel of at least one of the outer wing structure and the center wing structure gradually increases in a ramp region on at least one side of each padding area at the wing joint; and the stringer flange has a flange thickness that decreases in the ramp region as the panel thickness increases.

[0282] Clause A7: A wing assembly according to Clause A5 or A6, wherein at least one of the wing stringers of at least one of the outer wing structure and the center wing structure has its stringer flange connected along an adhesive line to a corresponding one of the upper skin panel and the lower skin panel; and the adhesive line terminates at a distance not exceeding 0.50 inches from the stringer end, and defines a flange-skin gap between the stringer flange and the skin panel.

[0283] Clause A8: A wing assembly according to any one of Clauses A5 to A7, wherein each outer wing structure is configured to include a fuel tank having an outer end of a box; and each outer wing structure has one or more outer wing stringers connected to at least one of an upper outer wing skin panel and a lower outer wing skin panel, and extending along the wingspan direction from the inner end of the outer wing and extending beyond the outer side of the outer end of the box.

[0284] Clause A9: The wing assembly according to any one of Clauses A1 to A8 further includes a center wing upper skin panel and a center wing lower skin panel connected to the center wing front spars and the center wing rear spars; an outer wing upper skin panel and an outer wing lower skin panel connected to the outer wing front spars and the outer wing rear spars; and wherein at least one of the outer wing structure and the center wing structure has one or more wing ribs, each wing rib being connected to the front spars, rear spars, upper skin panel, and lower skin panel of the outer wing structure and the center wing structure, respectively.

[0285] Clause A10: A method of manufacturing a wing assembly for an aircraft, comprising attaching the inner end of the outer wing of each of a pair of outer wing structures to the opposite outer end of a central wing structure, thereby defining a pair of wing joints for attaching the outer wing structures to the central wing structure; and wherein the central wing structure has a central wing forward sparb and a central wing aft sparb, each having a sparb end located at the outer end of each central wing, and an engine mounting position on each of the opposite sides of the wing centerline, and each of the opposite sparb ends of the central wing forward sparb and the central wing aft sparb being located at each wing joint no more inward than the engine centerline associated with the engine mounting position, and no more outward than 10 percent of the distance between the engine centerline and the wing centerline.

[0286] Clause A11: The method according to Clause A10 further includes connecting the outer wing upper skin panel and the outer wing lower skin panel to the outer wing front spars and outer wing rear spars of each outer wing structure; wherein at least one of the outer wing upper skin panel and the outer wing lower skin panel consists of an area layer having a constant panel thickness along the wingspan direction from the inner end of the outer wing to the wingtip; and the area layer includes one or more 0-degree layer strips interwoven within the area layer and extending in the wingspan direction.

[0287] Clause A12: The method described in Clause A10 or A11, wherein connecting each outer wing structure to the center wing structure includes connecting the center wing front spar to the outer wing front spar of the outer wing structure using a front spar splice plate; and connecting the center wing rear spar to the outer wing rear spar of the outer wing structure using a rear spar splice plate.

[0288] Clause A13: The method according to any one of Clauses A10 to A12, wherein connecting each outer wing structure to the center wing structure includes connecting the outer wing transverse rib on the inner end of the outer wing of the outer wing structure to the center wing transverse rib on the outer end of the center wing of the center wing structure, such that the outer wing transverse rib and the center wing transverse rib are arranged back to back.

[0289] Clause A14: The method according to any one of Clauses A10 to A13 further includes connecting one or more wing stringers to at least one of the upper and lower skin panels of at least one of the outer wing structure and the center wing structure, respectively; wherein each wing stringer has a stringer flange, at least one stringer web extending outwardly from the stringer flange, and a stringer transition at the stringer end, the height of the stringer web gradually decreasing along the stringer transition; and at least one of the upper and lower skin panels of the outer wing structure and the center wing structure has a partially raised area of ​​composite layer on the inner side of the skin panel at a chord position of the stringer flange of at least one of the wing stringers of at least one of the outer wing structure and the center wing structure, the partially raised area being located near the one or more stringer ends.

[0290] Clause A15: The method according to Clause A14, wherein at least one of the upper skin panel and the lower skin panel of at least one of the outer wing structure and the center wing structure is laid such that the panel thickness gradually increases in the ramp region on at least one side of each padding area at the wing joint; and at least one of the wing stringers of at least one of the outer wing structure and the center wing structure is laid such that the stringer flange has a flange thickness that decreases in the ramp region as the panel thickness increases.

[0291] Clause A16: The method according to Clause A14 or A15, wherein attaching one or more wing stringers to at least one skin panel comprises attaching a stringer flange of at least one of the wing stringers of at least one of the outer wing structure and the center wing structure to a corresponding one of the upper skin panel and the lower skin panel via an adhesive layer; terminating the adhesive layer at a distance of no more than 0.50 inches from the stringer end to define a flange-skin gap between the stringer flange and the skin panel; and filling the flange-skin gap with a non-adhesive gap filler.

[0292] Clause A17: The method according to any one of Clauses A14 to A16, wherein each of the outer wing structures is configured to include a fuel tank having an outer end of the box, and the step of connecting one or more wing stringers to at least one skin panel includes connecting one or more outer wing stringers to at least one of the outer wing upper skin panel and the outer wing lower skin panel in such a way that the outer wing stringer extends from the inner end of the outer wing along the wingspan direction and does not extend beyond the outer side of the outer end of the box.

[0293] Clause A18: The method according to any one of Clauses A10 to A17 further includes connecting one or more wing ribs to at least one of the outer wing structure and the center wing structure, including a front wing sparb and a rear wing sparb; and connecting an upper skin panel and a lower skin panel to at least one of the wing ribs, front wing sparb, and rear wing sparb of the outer wing structure and the center wing structure.

[0294] Clause A19: A method of loading an aircraft wing assembly, comprising providing a wing assembly having a pair of outer wing structures each connected to a central wing structure, the central wing structure including a pair of outer ends of a central wing and a pair of engine mounting locations located on opposite sides of the wing centerline; a central wing forward sparb and a central wing aft sparb, each having a sparb termination at each outer end of the central wing; each of the outer wing structures comprising an outer wing forward sparb and an outer wing aft sparb connected to the central wing forward sparb and the central wing aft sparb, respectively, to define a wing joint connecting the outer wing structures to the central wing structure; wherein the sparb terminations of the central wing forward sparb and the central wing aft sparb are located at each wing joint no more inward than the engine centerline associated with the engine mounting locations and no more outward than 10 percent of the distance between the engine centerline and the wing centerline; placing the wing assembly under static ground load conditions; and placing the wing assembly under dynamic load conditions.

[0295] Clause A20: The method described in Clause A19, wherein the ground static load condition relates to the aircraft being substantially stationary and under ground static load, under which the wing assembly is subjected to gravity due to the structural mass of the wing assembly; and the dynamic load condition relates to the aircraft motion and the wing assembly being subjected to at least one of the following: gravity due to the structural mass of the wing assembly; inertial forces due to the structural mass of the wing assembly; aerodynamic loads on the wing assembly; and control surface loads on the wing assembly.

[0296] Clause B1: A composite wing assembly for an aircraft, comprising a center wing structure including: a pair of outer ends of the center wing and a pair of engine mounting positions located on opposite sides of the wing centerline; a front sparb and a rear sparb of the center wing, each having a sparb end at the outer end of the center wing; a pair of outer wing structures, each outer wing structure including a front sparb and a rear sparb of the outer wing, configured to be respectively connected to the front sparb and the rear sparb of the center wing to define wing joints connecting the outer wing structures to the center wing structure; an upper skin panel and a lower skin panel of the outer wing, each having a shape along the wingspan. The structure comprises an area layer of constant panel thickness and is connected to the outer wing front spars and outer wing rear spars; one or more slats, which are joined to the inner surface of at least one of the upper and lower skin panels and extend in the wingspan direction, each slat consisting of a stack of slat layers narrower than the area layer; and wherein the central wing structure is configured such that the spars ends of the central wing front spars and central wing rear spars are located at each wing joint no more inward than the engine centerline associated with the engine mounting location, and no more outward than 10 percent of the distance between the engine centerline and the wing centerline.

[0297] Clause B2: The wing assembly according to Clause B1, wherein each slat has a slat centerline extending along the longitudinal direction of the slat; and the slats are spaced apart from each other at a chordal spacing of 6-11 inches between the slat centerlines.

[0298] Clause B3: A wing assembly as described in Clause B1 or B2, wherein at least some of the slats have a slat width between 3 and 6 inches.

[0299] Clause B4: A wing assembly according to any one of Clauses B1 to B3, wherein each slat consists of 10 to 70 slat layers; and the slat layers include 0-degree layers and non-0-degree layers, wherein the 0-degree layers comprise the maximum percentage of slat layers in each slat.

[0300] Clause B5: A wing assembly according to any one of Clauses B1 to B4, wherein at least some of the slats have a slat thickness that gradually decreases in the wingspan direction.

[0301] Clause B6: A wing assembly according to any one of Clauses B1 to B5, wherein the slat bending stiffness of each slat is approximately 40-60% of the combined slat-skin portion bending stiffness of the slat and the skin panel portion of the upper or lower skin panel supported by the slat; and the skin panel portion has a panel portion width defined by one of the following: a panel portion width extending between midpoints located on opposite sides of the slat, each midpoint being located at the midpoint between the slat and the adjacent slat; or a panel portion width extending between a sparb on one side of the slat and the midpoint between the slat and the adjacent slat on the opposite side of the slat.

[0302] Clause B7: A wing assembly according to any one of Clauses B1 to B6, wherein at least one of the upper skin panel and the lower skin panel comprises a 0-degree layer and a 90-degree layer; and the ratio of the 0-degree layer to the 90-degree layer is between 1.75 and 2.5.

[0303] Clause B8: A wing assembly according to any one of Clauses B1 to B7, wherein each wing joint includes a front spar splice plate connecting the center wing front spar to the outer wing front spar; and a rear spar splice plate connecting the center wing rear spar to the outer wing rear spar.

[0304] Clause B9: A wing assembly according to any one of Clauses B1 to B8, wherein the center wing structure includes a center wing transverse rib located at the outer end of each center wing; each outer wing structure includes an outer wing transverse rib located at the inner end of the outer wing; and at each wing joint, the outer wing transverse rib and the center wing transverse rib are arranged back-to-back and mechanically fastened to each other.

[0305] Clause B10: A method of manufacturing a composite wing assembly for an aircraft, comprising attaching one or more slats to the inner surface of at least one of an outer wing upper skin panel and an outer wing lower skin panel in each of an outer wing structure, each slat extending in the spanwise direction and comprising a stack of slat layers, each of the outer wing upper skin panel and the outer wing lower skin panel consisting of an area layer having a constant panel thickness along the spanwise direction, the slat layer having a narrower width than the area layer; attaching the outer wing upper skin panel and the outer wing lower skin panel to the outer wing forward spars and outer wing rear spars of the outer wing structure; and attaching each outer wing... The outer wing forward spars and outer wing rear spars of the wing structure are respectively connected to the center wing forward spars and center wing rear spars of the center wing structure to define a pair of wing joints at the opposite outer ends of the center wing; and wherein the center wing structure has engine mounting positions on each of the opposite sides of the wing centerline, and the spars ends of the center wing forward spars and center wing rear spars are located at each wing joint at a position no more inward than the engine centerline associated with the engine mounting position, and no more outward than 10 percent of the distance between the engine centerline and the wing centerline.

[0306] Clause B11: The method of Clause B10, wherein, prior to attaching the slats to the skin panel, the method includes positioning the slats on the skin panel at chordal spacing of 6-11 inches between the slat centerlines.

[0307] Clause B12: The method described in Clause B10 or B11 further includes laying multiple layers of slats such that at least some of the slats have a slat width between 3 and 6 inches.

[0308] Clause B13: The method according to any one of Clauses B10 to B12 further includes laying a plurality of 0-degree layers and non-0-degree layers to produce slats, wherein the 0-degree layers comprise a maximum percentage of the slat layers in at least one slat.

[0309] Clause B14: The method according to any one of Clauses B10 to B13 further includes laying multiple slat layers, and including at least one ply reduction in one or more slat layers for reducing slat stiffness in the spanwise direction.

[0310] Clause B15: The method according to any one of Clauses B10 to B14 further includes laying each slat to have a slat bending stiffness of approximately 40-60% of the combined slat-skin portion bending stiffness of the slat and skin panel portion; the skin panel portion having a panel portion width defined by one of the following: a panel portion width extending between midpoints located on opposite sides of the slat, each midpoint being located at the midpoint between the slat and the adjacent slat; or a panel portion width extending between a spar on one side of the slat and the midpoint between the slat and the adjacent slat on the opposite side of the slat.

[0311] Clause B16: The method according to any one of Clauses B10 to B15 further includes laying multiple area layers to produce an upper or lower outer wing skin panel, the area layers comprising a ratio of 0-degree to 90-degree layers between 1.75 and 2.5.

[0312] Clause B17: The method according to any one of Clauses B10 to B16, wherein connecting each outer wing structure to the center wing structure comprises connecting the center wing front spar to the outer wing front spar using a front spar splice plate; and connecting the center wing rear spar to the outer wing rear spar using a rear spar splice plate.

[0313] Clause B18: The method according to any one of Clauses B10 to B17, wherein connecting each outer wing structure to the center wing structure includes connecting the outer wing transverse rib on the inner end of the outer wing of the outer wing structure to the center wing transverse rib on the outer end of the center wing of the center wing structure, such that the outer wing transverse rib and the center wing transverse rib are arranged back to back.

[0314] Clause B19: A method of loading a composite wing assembly for an aircraft, comprising providing a wing assembly having a pair of outer wing structures each connected to a central wing structure, the central wing structure including a pair of outer wing ends and a pair of engine mounting positions located on opposite sides of the wing centerline; a central wing forward sparb and a central wing aft sparb, each having a sparb end at each outer wing end; each of the outer wing structures comprising an outer wing forward sparb and an outer wing aft sparb configured to be connected to the central wing forward sparb and central wing aft sparb respectively to define a wing joint, the wing joint connecting the outer wing structures to the central wing structure at the outer wing ends; an outer wing upper skin panel and an outer wing lower skin panel, each having a free edge along... The wing assembly comprises an area layer of panel thickness with a constant thickness in the wingspan direction, and is connected to the outer wing's front and rear spars; one or more slats, which are attached to the inner surface of at least one of the upper and lower skin panels and extend in the wingspan direction, each slat consisting of a stack of slat layers narrower than the area layer; wherein the center wing structure is configured such that the spars ends of the center wing's front and rear spars are located at each wing joint no more inward than the engine centerline associated with the engine mounting location, and no more outward than 10 percent of the distance between the engine centerline and the wing centerline; the wing assembly is subjected to static ground load conditions; and the wing assembly is subjected to dynamic load conditions.

[0315] Clause B20: The method described in Clause B19, wherein the ground static load condition relates to the aircraft being substantially stationary and under ground static load, under which the wing assembly is subjected to gravity due to the structural mass of the wing assembly; and

[0316] Dynamic load conditions are related to aircraft motion and at least one of the following: gravity due to the structural mass of the wing assembly; inertial forces due to the structural mass of the wing assembly; aerodynamic loads on the wing assembly; and control surface loads on the wing assembly.

[0317] Clause C1: A composite wing assembly for an aircraft, comprising: a center wing structure including a pair of outer wing ends and a pair of engine mounting positions located on opposite sides of the wing centerline; a center wing forward sparb and a center wing aft sparb, each having a sparb termination at each outer wing end; a pair of outer wing structures, each outer wing structure including an outer wing forward sparb and an outer wing aft sparb configured to be respectively connected to the center wing forward sparb and the center wing aft sparb to define wing joints connecting the outer wing structures to the center wing structure; an outer wing upper skin panel and an outer wing lower skin panel, each composed of an area layer having a constant panel thickness along the wingspan direction, and It is connected to the front wing spars and the rear wing spars of the outer wing; one or more outer wing stringers, which are connected to at least one of the upper skin panels and the lower skin panels of the outer wing and extend in the spanwise direction, and each has a cap-shaped section with a stringer cover, the stringer cover containing one or more 0-degree layer strips interwoven within the stringer cover; and wherein the center wing structure is configured such that the wing spars ends of the front wing spars and the rear wing spars of the center wing are located at each wing joint no more inward than the engine centerline associated with the engine mounting position, and no more outward than 10 percent of the distance between the engine centerline and the wing centerline.

[0318] Clause C2: The wing assembly as described in Clause C1, wherein the 0-degree layer strip in the stringer cover has one or more ply reductions to reduce the bending stiffness of the outer wing structure along the span direction.

[0319] Clause C3: The wing assembly according to Clause C1 or C2, wherein each outer wing stringer further includes a stringer base laminate positioned between the stringer flange of each outer wing stringer and the upper skin panel and / or the lower skin panel of the outer wing; and one or more 0-degree layer strips interwoven within the stringer layer of the stringer base laminate.

[0320] Clause C4: The wing assembly according to Clause C3, wherein the one or more 0-degree layer strips interwoven in the stringer base laminate have one or more ply reductions along the spanwise direction to reduce the bending stiffness of the outer wing structure at each spanwise location.

[0321] Clause C5: A wing assembly according to any one of Clauses C1 to C4, wherein the stringer bending stiffness of each outer wing stringer is approximately 40-60% of the combined stringer-skin portion bending stiffness of the outer wing stringer and the skin panel portion of the upper or lower skin panel supported by the outer wing stringer upper or lower skin panel; and the skin panel portion has a panel portion width defined by one of the following: a panel portion width extending between midpoints located on opposite sides of the outer wing stringer, each midpoint being located at an intermediate position between the outer wing stringer and the adjacent outer wing stringer; or a panel portion width extending between a sparse on one side of the outer wing stringer and the midpoint between the outer wing stringer and the adjacent outer wing stringer on the opposite side of the outer wing stringer.

[0322] Clause C6: A wing assembly according to any one of Clauses C1 to C5, wherein at least one of the outer wing upper skin panel and the outer wing lower skin panel has one or more 0-degree layer strips that extend in the span direction and are interwoven within the area layer at one or more stringer mounting locations of the outer wing stringers.

[0323] Clause C7: A wing assembly according to any one of Clauses C1 to C6, wherein the 0-degree layer strip in the skin panel has one or more ply reductions positioned according to the desired bending stiffness of the outer wing structure.

[0324] Clause C8: A wing assembly according to any one of Clauses C1-C7, wherein each wing joint includes a front spar splice plate connecting the center wing front spar to the outer wing front spar; and a rear spar splice plate connecting the center wing rear spar to the outer wing rear spar.

[0325] Clause C9: A wing assembly according to any one of Clauses C1 to C8, wherein the center wing structure includes a center wing transverse rib located at the outer end of the center wing; each outer wing structure includes an outer wing transverse rib located at the inner end of the outer wing; and at each wing joint, the outer wing transverse rib and the center wing transverse rib are arranged back-to-back and mechanically fastened to each other.

[0326] Clause C10: A method of manufacturing a composite wing assembly for an aircraft, comprising attaching one or more outer wing stringers to at least one of an upper outer wing skin panel and an outer wing lower outer wing skin panel, each outer wing stringer extending in the spanwise direction and consisting of a plurality of stringer layers, and having a stringer cap including one or more 0-degree layer strips, each of the upper outer wing skin panel and the lower outer wing skin panel consisting of an area layer having a panel thickness constant along the spanwise direction; and attaching the upper outer wing skin panel and the lower outer wing skin panel to the outer wing front spars and the outer wing rear spars of the outer wing structure. A beam; connecting the inner end of the outer wing of each of a pair of outer wing structures to the opposite outer end of the center wing of a central wing structure, thereby defining a pair of wing joints; and wherein the central wing structure has an engine mounting position on each of the opposite sides of the wing centerline, and the sparsity ends of the front and rear spars of the central wing are located at each wing joint no more inward than the engine centerline associated with the engine mounting position, and no more outward than 10 percent of the distance between the engine centerline and the wing centerline.

[0327] Clause C11: The method described in Clause C10 further includes laying multiple stringer layers to produce an outer wing stringer; interlacing the one or more 0-degree layer strips within the stringer layers of the stringer cap of the wing stringer; and incorporating one or more ply reductions in the 0-degree layer strips of the stringer cap to reduce the bending stiffness of the outer wing structure along the span direction.

[0328] Clause C12: The method according to Clause C10 or C11, wherein attaching the outer wing stringers to the outer wing upper skin panel and the outer wing lower skin panel includes attaching a stringer base laminate to the outer wing upper skin panel and / or the outer wing lower skin panel, such that the stringer flange of each outer wing stringer is mounted on the stringer base laminate; and the stringer base laminate includes one or more 0-degree layer strips interwoven within the stringer layers of the stringer base laminate.

[0329] Clause C13: The method described in Clause C12 further includes laying each stringer base laminate using stringer layers; interlacing the one or more 0-degree layer strips within the stringer layers of the stringer base laminate; and incorporating one or more ply reductions in the 0-degree layer strips of the stringer base laminate to reduce the bending stiffness of the outer wing structure along the span direction.

[0330] Clause C14: The method according to any one of Clauses C10 to C13, wherein each outer wing stringer is laid to have a stringer bending stiffness of approximately 40-60% of the combined slat-skin portion bending stiffness of the slat and the skin panel portion of the outer wing upper skin panel or the outer wing lower skin panel; the skin panel portion has a panel portion width defined by one of the following: a panel portion width extending between midpoints located on opposite sides of the outer wing stringer, each midpoint being located at the midpoint between the outer wing stringer and the adjacent outer wing stringer; or a panel portion width extending between a spar on one side of the outer wing stringer and the midpoint between the outer wing stringer and the adjacent outer wing stringer on the opposite side of the outer wing stringer.

[0331] Clause C15: The method according to any one of Clauses C10 to C14, wherein a plurality of area layers are laid to produce an upper outer wing skin panel and a lower outer wing skin panel; at one or more stringer mounting locations of the outer wing stringers on the upper outer wing skin panel and the lower outer wing skin panel, one or more 0-degree layer strips are interwoven within the area layers.

[0332] Clause C16: The method according to any one of Clauses C10 to C15 further includes incorporating one or more ply reductions in the 0-degree layer strips of the outer wing upper skin panel or the outer wing lower skin panel to reduce the bending stiffness of the outer wing structure along the wingspan direction.

[0333] Clause C17: The method according to any one of Clauses C10 to C16, wherein connecting each outer wing structure to the center wing structure comprises connecting the center wing front spar to the outer wing front spar using a front spar splice plate; and connecting the center wing rear spar to the outer wing rear spar using a rear spar splice plate.

[0334] Clause C18: The method according to any one of Clauses C10 to C17, wherein connecting each outer wing structure to the center wing structure includes connecting the outer wing transverse rib on the inner end of the outer wing of the outer wing structure to the center wing transverse rib on the outer end of the center wing of the center wing structure, such that the outer wing transverse rib and the center wing transverse rib are arranged back to back.

[0335] Clause C19: A method of loading a composite wing assembly for an aircraft, comprising providing a wing assembly having a pair of outer wing structures each connected to a central wing structure, the central wing structure including a pair of outer ends of a central wing and a pair of engine mounting positions located on opposite sides of the wing centerline; a central wing front sparb and a central wing rear sparb, each having a sparb end at the outer end of the central wing; each of the outer wing structures comprising an outer wing front sparb and an outer wing rear sparb configured to be connected to the central wing front sparb and the central wing rear sparb respectively to define a wing joint connecting the outer wing structures to the central wing structure; an outer wing upper skin panel and an outer wing lower skin panel, each consisting of an area layer having a constant panel thickness along the wingspan, and connected... The outer wing includes a front spar and a rear spar; one or more outer wing stringers, which are connected to at least one of the upper and lower skin panels of the outer wing and extend in the spanwise direction, and each has a cap-shaped section with a stringer cap containing one or more 0-degree layer strips interwoven within the stringer cap; wherein the center wing structure is configured such that the spar ends of the front and rear spars of the center wing are located at each wing joint no more inward than the engine centerline associated with the engine mounting location, and no more outward than 10 percent of the distance between the engine centerline and the wing centerline; the wing assembly is placed under static ground load conditions; and the wing assembly is placed under dynamic load conditions.

[0336] Clause C20: The method described in Clause C19, wherein the ground static load condition relates to the aircraft being substantially stationary and under ground static load, under which the wing assembly is subjected to gravity due to the structural mass of the wing assembly; and the dynamic load condition relates to the motion of the aircraft and to at least one of the following: gravity due to the structural mass of the wing assembly; inertial forces due to the structural mass of the wing assembly; aerodynamic loads on the wing assembly; and control surface loads on the wing assembly.

[0337] Clause D1: A composite wing assembly for an aircraft, comprising: a center wing structure including a pair of outer ends of the center wing and a pair of engine mounting positions located on opposite sides of the wing centerline; a center wing forward sparb and a center wing aft sparb, each having a sparb termination at the outer ends of the center wing; a pair of outer wing structures, each outer wing structure including an outer wing forward sparb and an outer wing aft sparb configured to be respectively connected to the center wing forward sparb and the center wing aft sparb to define a wing joint connecting the outer wing structures to the center wing structure; an outer wing upper panel assembly and an outer wing lower panel assembly, each extending along the wingspan and connected to the outer wing forward sparb. The wing spars and outer wing rear spars, each panel assembly including an outer skin panel composed of area layers and having a constant panel thickness along the wingspan; a hat-shaped section panel attached to the inner side of the outer skin panel and having multiple generally parallel hat-shaped sections, each extending in the wingspan; and wherein the center wing structure is configured such that the spars ends of the center wing forward spars and center wing rear spars are located at each wing joint no more inward than the engine centerline associated with the engine mounting location, and no more outward than 10 percent of the distance between the engine centerline and the wing centerline.

[0338] Clause D2: The wing assembly according to Clause D1, wherein the cap section panel is composed of a plurality of cap section panel layers; each cap section includes a pair of cap section flanges, a pair of cap section webs extending outward from the cap section flanges, and cap section caps interconnected with the cap section webs; and each cap section cap includes one or more 0-degree layer strips interwoven within the cap section panel layers of the cap section panel.

[0339] Clause D3: The wing assembly according to Clause D2, wherein at least one 0-degree layer strip in the cap-shaped section has a ply reduction to reduce the bending stiffness of the outer wing structure along the span direction.

[0340] Clause D4: A wing assembly according to any one of Clauses D1 to D3, wherein each cap section has a cap section centerline extending along the longitudinal direction of the cap section; and the cap sections are spaced apart from each other by a chordal spacing of 6-11 inches between the cap section centerlines.

[0341] Clause D5: A wing assembly according to any one of Clauses D1 to D4, wherein each panel assembly comprises a plurality of slats extending in the spanwise direction between the plurality of cap-shaped sections and captured between the outer skin panel and the cap-shaped section panel, each slat comprising a plurality of slat layers including one or more 0-degree layer strips for increasing the bending stiffness of the outer skin panel.

[0342] Clause D6: The wing assembly according to Clause D5, wherein the one or more 0-degree layer strips interwoven in the slats have one or more ply reductions along the spanwise direction to reduce the bending stiffness of the outer wing structure at each spanwise location.

[0343] Clause D7: A wing assembly according to any one of Clauses D1 to D6, wherein a plurality of cap-shaped section cavities are defined between the cap-shaped section and the outer skin panel, respectively; and each panel assembly includes a plurality of foam members occupying the plurality of cap-shaped section cavities.

[0344] Clause D8: A wing assembly according to any one of Clauses D1 to D7, wherein each wing joint includes a front spar splice plate connecting the center wing front spar to the outer wing front spar; and a rear spar splice plate connecting the center wing rear spar to the outer wing rear spar.

[0345] Clause D9: A wing assembly according to any one of Clauses D1 to D8, wherein the center wing structure includes a center wing transverse rib located at the outer end of each center wing; each outer wing structure includes an outer wing transverse rib located at the inner end of the outer wing; and at each wing joint, the outer wing transverse rib and the center wing transverse rib are arranged back-to-back and mechanically fastened to each other.

[0346] Clause D10: A method of manufacturing a composite wing assembly for an aircraft, comprising attaching a cap-section panel to an outer skin panel of at least one of an upper outer wing panel assembly and an lower outer wing panel assembly, the cap-section panel having a plurality of generally parallel cap-sections, each cap-section extending in the spanwise direction, the outer skin panel consisting of an area layer having a panel thickness having a constant thickness along the spanwise direction; connecting the upper outer wing panel assembly and the lower outer wing panel assembly to an outer wing forward sparb and an outer wing aft sparb of an outer wing structure; connecting the inner outer ends of each of a pair of outer wing structures to opposite outer central wing ends of a central wing structure to define a pair of wing joints; and wherein the central wing structure has engine mounting positions on each of opposite sides of the wing centerline, and the sparsity ends of the central wing forward sparb and the central wing aft sparb are located at each wing joint no more inward than the engine centerline associated with the engine mounting position, and no more outward than 10 percent of the distance between the engine centerline and the wing centerline.

[0347] Clause D11: The method according to Clause D10 further includes laying a plurality of cap-shaped section panel layers to produce a cap-shaped section panel; and at each of the respective plurality of cap-shaped sections, interlacing one or more 0-degree layer strips within the cap-shaped section panel layer.

[0348] Clause D12: The method according to Clause D11, wherein at least one of the 0-degree layer strips of the cap-shaped section cover includes a ply reduction to reduce the bending stiffness of the outer wing structure along the span direction.

[0349] Clause D13: The method according to any one of Clauses D10 to D12, wherein laying the cap-shaped panel layer to produce the cap-shaped panel comprises spacing the cap-shaped sections apart from each other at a chordal spacing of 6-11 inches between the centerlines of the cap-shaped sections.

[0350] Clause D14: The method according to any one of Clauses D10 to D13 further includes laying a plurality of slat layers on an outer skin panel to produce a plurality of slats, each slat extending in the spanwise direction and located between the plurality of cap-shaped sections; and attaching a cap-shaped section panel to the slats and the outer skin panel such that the slats are trapped between the outer skin panel and the cap-shaped section panel.

[0351] Clause D15: The method according to Clause D14, wherein the slatted laminate includes one or more 0-degree layers, which include layup reduction in the one or more 0-degree layers to reduce the bending stiffness of the outer wing structure at each spanwise location.

[0352] Clause D16: The method according to Clause D14 or D15, wherein laying the cap-shaped section panel includes laying the cap-shaped section panel on a plurality of foam members to create the plurality of cap-shaped sections, each cap-shaped section having a pair of cap-shaped section flanges, a pair of cap-shaped section webs extending outwardly from the cap-shaped section flanges, a cap-shaped section cap connecting the cap-shaped section webs to each other, and a plurality of cap-shaped section cavities respectively defined between the plurality of cap-shaped sections and the outer skin panel; and attaching the cap-shaped section panel to the outer skin panel such that the foam members are trapped therebetween.

[0353] Clause D17: The method according to any one of Clauses D10 to D16, wherein connecting each outer wing structure to the center wing structure comprises connecting the center wing front spar to the outer wing front spar using a front spar splice plate; and connecting the center wing rear spar to the outer wing rear spar using a rear spar splice plate.

[0354] Clause D18: The method according to any one of Clauses D10 to D17, wherein connecting each outer wing structure to the center wing structure includes connecting the outer wing transverse rib on the inner end of the outer wing of the outer wing structure to the center wing transverse rib on the outer end of the center wing of the center wing structure, such that the outer wing transverse rib and the center wing transverse rib are arranged back to back.

[0355] Clause D19: A method of loading a composite wing assembly for an aircraft, comprising providing a wing assembly having a pair of outer wing structures each connected to a central wing structure, the central wing structure including: a pair of outer ends of a central wing and a pair of engine mounting locations located on opposite sides of the wing centerline; a central wing forward sparb and a central wing aft sparb, each having a sparb termination at each outer end of the central wing; each outer wing structure including an outer wing forward sparb and an outer wing aft sparb configured to be connected to the central wing forward sparb and central wing aft sparb respectively to define a wing joint connecting the outer wing structure to the central wing structure; an outer wing upper panel assembly and an outer wing lower panel assembly, each extending along the wingspan and connected to the outer wing forward sparb. The wing assembly comprises a spars and an outer wing aft spars, each panel assembly including an outer skin panel consisting of area layers and having a constant panel thickness along the wingspan; a cap-shaped section panel connected to the outer skin panel and having multiple generally parallel cap-shaped sections, each extending in the wingspan; wherein the center wing structure is configured such that the spars ends of the center wing forward spars and center wing aft spars are located at each wing joint no more inward than the engine centerline associated with the engine mounting location, and no more outward than 10 percent of the distance between the engine centerline and the wing centerline; and the wing assembly is placed under static ground load conditions; and the wing assembly is placed under dynamic load conditions.

[0356] Clause D20: The method according to Clause D19, wherein the ground static load condition relates to the aircraft being substantially stationary and under ground static load, under which the wing assembly is subjected to gravity due to the structural mass of the wing assembly; and the dynamic load condition relates to the motion of the aircraft and to at least one of the following: gravity due to the structural mass of the wing assembly; inertial forces due to the structural mass of the wing assembly; aerodynamic loads on the wing assembly; and control surface loads on the wing assembly.

[0357] Further modifications and improvements to this disclosure will likely be apparent to those skilled in the art. Therefore, the specific combinations of components described and illustrated herein are intended only to illustrate certain embodiments of this disclosure and are not intended to be used as limitations on alternative embodiments or apparatuses within the spirit and scope of this disclosure.

Claims

1. A wing assembly (120) for an aircraft (100), comprising: The center wing structure (150) includes: a pair of outer ends (168) of the center wing and a pair of engine mounting positions located on opposite sides of the wing centerline (124); a front sparsity (152) and a rear sparsity (154) of the center wing, each having a sparsity end (156) at each outer end (168) of the center wing; and a transverse rib (170) of the center wing at each outer end (168) of the center wing. A pair of outer wing structures (210), each outer wing structure including: a front wing sparb (212); a rear wing sparb (214); and a transverse wing rib (232) at the inner end (230) of the outer wing. The outer wing structure (210) is generally flat and straight, and has a smaller bending stiffness than the central wing structure (150). The outer wing front spar (212) and the outer wing rear spar (214) are configured to be connected to the central wing front spar (152) and the central wing rear spar (154) respectively, so as to define the wing joint (172) that connects the outer wing structure (210) to the central wing structure (150). At each wing joint (172), the outer wing transverse rib (232) and the center wing transverse rib (170) are configured to be positioned back-to-back and mechanically fastened to each other. The central wing transverse rib (170) and the outer wing transverse rib (232) are configured to provide a load path across the wing joint (172) from the outer wing structure (210) to the central wing structure (150), and to accommodate the difference in bending stiffness between the outer wing structure (210) and the central wing structure (150); and The central wing structure (150) is configured such that the spar ends (156) of the central wing front spar (152) and the central wing rear spar (154) are located at each wing joint (172) no more inward than the engine centerline (106) associated with the engine mounting position, and no more outward than 10 percent of the distance between the engine centerline (106) and the wing centerline (124).

2. The wing assembly (120) according to claim 1 further comprises: The upper skin panel (240) and the lower skin panel (242) of the outer wing are connected to the front wing spars (212) and the rear wing spars (214) of the outer wing. At least one of the outer wing upper skin panel (240) and the outer wing lower skin panel (242) is composed of an area layer (302) having a constant panel thickness (194) along the wingspan direction from the inner end (230) of the outer wing towards the wingtip (126); and At least one of the outer wing upper skin panel (240) and the outer wing lower skin panel (242) has one or more 0-degree stripe groups (250), each 0-degree stripe group extending in the wingspan direction, and each 0-degree stripe group includes one or more 0-degree layer stripes (252) interwoven within the area layer (302).

3. The wing assembly (120) according to claim 1, wherein, Each wing joint (172) includes: A front wing sparsity splice (216) connects the center wing front wing spars (152) to the outer wing front wing spars (212); and The rear wing spars splice plate (218) connects the center wing rear wing spars (154) to the outer wing rear wing spars (214).

4. The wing assembly (120) according to claim 1 further comprises: The upper skin panel (190) and the lower skin panel (192) of the center wing are connected to the front wing spars (152) and the rear wing spars (154) of the center wing. The upper skin panel (240) and the lower skin panel (242) of the outer wing are connected to the front wing spars (212) and the rear wing spars (214) of the outer wing. Wherein, at least one of the outer wing structure (210) and the central wing structure (150) has one or more wing stringers (270, 272), each wing stringer being connected to one of the upper skin panel and the lower skin panel of at least one of the outer wing structure (210) and the central wing structure (150); Each of the wing stringers (270, 272) has a stringer flange (276), at least one stringer web (278) extending outward from the stringer flange (276), and a stringer transition section (280) at the end of the stringer, the stringer web (278) gradually decreasing in height along the stringer transition section (280); and At least one of the upper and lower skin panels of the outer wing structure (210) and the center wing structure (150) has a partially raised area (200) of a composite layer (300) located on the inside of the skin panel (190, 192, 240, 242) at the chord position of the string flange (276) of at least one of the wing stringers (270, 272) of the outer wing structure (210) and the center wing structure (150), the partially raised area (200) being located near one or more stringer ends.

5. The wing assembly (120) according to claim 4, wherein: The panel thickness (194) of at least one of the upper and lower skin panels of at least one of the outer wing structure (210) and the central wing structure (150) gradually increases within a ramp region (202) on at least one side of each locally raised area (200) located at the wing joint (172); and The stringer flange (276) has a flange thickness (164) that decreases within the slope region (202) as the panel thickness (194) increases.

6. The wing assembly (120) according to claim 4, wherein: The stringer flange (276) of at least one of the wing stringers (270, 272) of at least one of the outer wing structure (210) and the center wing structure (150) is connected along the bonding line (282) to a corresponding one of the upper skin panel (190, 240) and the lower skin panel (192, 242); and The adhesive line (282) terminates at a distance of no more than 0.50 inches from the end of the stringer and defines a flange-skin gap (286) between the stringer flange (276) and the skin panel (190, 192, 240, 242).

7. The wing assembly (120) according to claim 4, wherein: Each of the outer wing structures (210) comprises a fuel tank (128) having an outer end (134); and Each outer wing structure (210) has one or more outer wing stringers (272) connected to at least one of the outer wing upper skin panel (240) and the outer wing lower skin panel (242), and extending along the wingspan direction from the inner end (230) of the outer wing and not extending further outward than the outer end (134) of the box.

8. The wing assembly (120) according to any one of claims 1 to 7, further comprising: The upper skin panel (190) and the lower skin panel (192) of the center wing are connected to the front wing spars (152) and the rear wing spars (154) of the center wing. The upper skin panel (240) and the lower skin panel (242) of the outer wing are connected to the front wing spars (212) and the rear wing spars (214) of the outer wing; and At least one of the outer wing structure (210) and the center wing structure (150) has one or more wing ribs (174, 234), each wing rib being connected to the front wing spars (152, 212), rear wing spars (154, 214), upper skin panel (190, 240) and lower skin panel (192, 242) of the outer wing structure (210) and the center wing structure (150), respectively.

9. A method of manufacturing a wing assembly (120) for an aircraft (100), comprising: The inner end (230) of each of the outer wings in a pair of outer wing structures (210) is respectively connected to the opposite outer end (168) of the central wing of the central wing structure (150) to define a pair of wing joints (172) connecting the outer wing structure (210) to the central wing structure (150). Each of the outer wing structures (210) is generally flat and straight, and has a smaller bending stiffness than the central wing structure (150). The process of connecting each outer wing structure (210) to the central wing structure (150) includes connecting the outer wing transverse rib (232) on the inner end (230) of the outer wing of the outer wing structure (210) to the central wing transverse rib (170) on the outer end (168) of the central wing of the central wing structure (150), such that the outer wing transverse rib (232) and the central wing transverse rib (170) are arranged back to back. The central wing transverse rib (170) and the outer wing transverse rib (232) are configured to provide a load path across the wing joint (172) from the outer wing structure (210) to the central wing structure (150), and to accommodate the difference in bending stiffness between the outer wing structure (210) and the central wing structure (150); and The central wing structure (150) has a central wing front sparb (152) and a central wing rear sparb (154), each having a sparb end (156) located at the outer end (168) of each central wing. The central wing structure also has engine mounting positions on each opposite side of the wing centerline (124), and each of the opposite sparb ends (156) of the central wing front sparb (152) and the central wing rear sparb (154) is located at each wing joint (172) at a position no more inward than the engine centerline (106) associated with the engine mounting position, and no more outward than 10 percent of the distance between the engine centerline (106) and the wing centerline (124).

10. The method of claim 9, further comprising: The upper outer wing skin panel (240) and the lower outer wing skin panel (242) are connected to the front wing spars (212) and the rear wing spars (214) of each of the outer wing structures (210). At least one of the outer wing upper skin panel (240) and the outer wing lower skin panel (242) is composed of an area layer (302) having a panel thickness (194) that is constant along the wingspan direction from the inner end (230) of the outer wing to the wingtip (126); and The area layer (302) includes one or more 0-degree layer strips (252) interwoven within the area layer (302) and extending in the wingspan direction.

11. The method according to claim 9, wherein, Connecting each outer wing structure (210) to the central wing structure (150) includes: The center wing front spar (152) is connected to the outer wing front spar (212) of the outer wing structure (210) using a front spar splice plate (216); and The center wing rear spar (154) is connected to the outer wing rear spar (214) of the outer wing structure (210) using a rear spar splice plate (218).

12. The method according to claim 9, further comprising: One or more wing stringers (270, 272) are respectively connected to at least one of the upper skin panel and lower skin panel of at least one of the outer wing structure (210) and the central wing structure (150); Each of the wing stringers (270, 272) has a stringer flange (276), at least one stringer web (278) extending outward from the stringer flange (276), and a stringer transition section (280) at the end of the stringer, wherein the stringer web (278) gradually decreases in height along the stringer transition section (280); and At least one of the upper and lower skin panels of the outer wing structure (210) and the center wing structure (150) has a partially raised area (200) of a composite layer (300) located on the inside of the skin panel (190, 192, 240, 242) at the chord position of the string flange (276) of at least one of the wing stringers (270, 272) of the outer wing structure (210) and the center wing structure (150), the partially raised area (200) being located near one or more stringer ends.

13. The method according to claim 12, wherein: At least one of the upper and lower skin panels of at least one of the outer wing structure (210) and the central wing structure (150) is laid such that the panel thickness (194) gradually increases within a ramp region (202) on at least one side of each locally raised area (200) at the wing joint (172); and At least one wing stringer (270, 272) of at least one of the outer wing structure (210) and the central wing structure (150) is laid such that the stringer flange (276) has a flange thickness (164) that decreases within the ramp region (202) as the panel thickness (194) increases.

14. The method according to claim 12, wherein, Connecting one or more wing stringers (270, 272) to at least one of the skin panels (190, 192, 240, 242) includes: The stringer flange (276) of at least one of the outer wing structure (210) and the central wing structure (150) is connected via an adhesive layer (284) to a corresponding one of the upper skin panel (190, 240) and the lower skin panel (192, 242); The adhesive layer (284) terminates at a distance of no more than 0.50 inches from the end of the stringer to define a flange-skin gap (286) between the stringer flange (276) and the skin panels (240, 242); and The flange-skin gap (286) is filled with non-adhesive gap filler.

15. The method according to claim 12, wherein, Each configuration in the outer wing structure (210) includes a fuel tank (128) having an outer end (134), and the steps of connecting one or more wing stringers (270, 272) to at least one skin panel (190, 192, 240, 242) include: One or more outer wing stringers (272) are connected in a manner to at least one of the outer wing upper skin panel (240) and the outer wing lower skin panel (242) such that the outer wing stringers (272) extend from the inner end (230) of the outer wing along the wingspan direction and do not extend further outward than the outer end (134) of the box.

16. The method according to any one of claims 9 to 15, further comprising: One or more wing ribs (174, 234) are connected to the front wing spars (152, 212) and rear wing spars (154, 214) of at least one of the outer wing structure (210) and the center wing structure (150). as well as The upper skin panel (190, 240) and the lower skin panel (192, 242) are connected to at least one of the wing ribs (174, 234), the front wing spars (152, 212) and the rear wing spars (154, 214) of at least one of the outer wing structure (210) and the center wing structure (150).

17. A method of loading a wing assembly (120) of an aircraft (100), comprising: A wing assembly (120) is provided, the wing assembly having an outer wing structure (210), each of the outer wing structures being connected to a central wing structure (150), the central wing structure (150) comprising: The engine mounting positions are located at the outer ends (168) of a pair of central wings and on opposite sides of the wing centerline (124); The center wing forward sparb (152) and center wing rear sparb (154) each have a sparb end (156) at the outer end (168) of each center wing. Central wing transverse rib (170) at the outer end (168) of each of the central wings; Each of the outer wing structures (210) includes: The outer wing front spars (212); the outer wing rear spars (214); and the outer wing transverse rib (232) located at the inner end (230) of the outer wing. The outer wing structure (210) is generally flat and straight, and has a smaller bending stiffness than the central wing structure (150). The outer wing front spar (212) and the outer wing rear spar (214) are configured to be connected to the central wing front spar (152) and the central wing rear spar (154) respectively, so as to define the wing joint (172) that connects the outer wing structure (210) to the central wing structure (150). At each wing joint (172), the outer wing transverse rib (232) and the center wing transverse rib (170) are configured to be positioned back-to-back and mechanically fastened to each other. The central wing transverse rib (170) and the outer wing transverse rib (232) are configured to provide a load path across the wing joint (172) from the outer wing structure (210) to the central wing structure (150), and to accommodate the difference in bending stiffness between the outer wing structure (210) and the central wing structure (150); and The spar ends (156) of the front wing spars (152) and the rear wing spars (154) of the center wing are located at each wing joint (172) at a position no more inward than the engine centerline (106) associated with the engine mounting position, and at a position no more outward than 10 percent of the distance between the engine centerline (106) and the wing centerline (124). The wing assembly (120) is placed under static ground load conditions; and The wing assembly (120) is placed under dynamic load conditions.

18. The method of claim 17, wherein: The ground static load conditions relate to the aircraft (100) that is substantially stationary and under ground static load, under which the wing assembly (120) is subjected to gravity due to the structural mass of the wing assembly (120); and The dynamic load conditions relate to the moving aircraft (100) and the wing assembly (120) subjected to at least one of the following: Gravity generated due to the structural mass of the wing assembly (120); Inertial forces generated due to the structural mass of the wing assembly (120); The aerodynamic loads on the wing assembly (120); and Control surface loads on the wing assembly (120).

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