Circumferential overlapping joint of the fuselage section

Abstract

Relates to a circumferential butt joint of the fuselage section. Specifically, a system and method for splicing a frame assembly are provided. One embodiment is a method for assembling a frame of an aircraft. The method includes the following steps: forming a first skin of a first circumferential section of the fuselage, the first skin including a distal portion that includes a lip and a shoulder; aligning a second skin of a second circumferential section of the fuselage with the shoulder such that the lip overlaps the second skin; and securing the first skin and the second skin together via a circumferential joint.

Description

Technical Field

[0001] The present disclosure relates to the field of assembly, and more particularly to the assembly of the airframe of an aircraft. Background Art

[0002] The airframe defines the mechanical structure of an aircraft. The airframe is made up of a plurality of components that provide the structural performance required for the aircraft. For example, a portion of the airframe for an aircraft fuselage may include a frame, a skin, and stringers that are mechanically coupled together (e.g., via co-bonding, co-curing, or fasteners) according to design parameters. As currently practiced, sections of the fuselage can be manufactured as full barrel sections, and these full barrel sections can be joined via circumferential splices. However, circumferential splicing involves adding splice straps and numerous fastening components around the entire perimeter of the fuselage section being joined. Thus, circumferential joints add a significant amount of weight to the aircraft and involve a great deal of labor in installing the circumferential joints, particularly with respect to aligning the splice straps with other fastening components.

[0003] Accordingly, there is a desire for a method and apparatus that addresses at least some of the problems discussed above, as well as other possible problems. Summary of the Invention

[0004] Embodiments described herein provide fuselage sections having skins that are sized to fit closely together (e.g., via shiplap joints, rabbets, or other features). These complementary skin sections enhance the strength of the circumferential joints between the barrel-shaped fuselage sections and can also eliminate the need for splice straps. This can reduce the amount of added weight involved in each circumferential joint and can additionally reduce the complexity of aligning and assembling each circumferential joint. Accordingly, the embodiments provided herein result in technical benefits in the form of increased (or equivalent) strength, reduced weight, and reduced labor as compared to other joints.

[0005] One embodiment is a method for assembling an airframe of an aircraft. The method includes the steps of forming a first skin of a first circumferential section of the fuselage. The first skin includes a distal portion that includes a lip and a shoulder. The method includes the steps of aligning a second skin of a second circumferential section of the fuselage with the shoulder such that the lip overlaps the second skin; and further includes the step of securing the first skin and the second skin together via a circumferential joint.

[0006] Another embodiment is a system including a portion of an airframe of an aircraft. The system includes a first skin of a first circumferential section of the fuselage, the first skin including a distal portion that includes a lip and a shoulder. The system further includes a second skin of a second circumferential section of the fuselage, the second skin being aligned with the shoulder such that the lip overlaps the second skin; and further includes a circumferential joint that secures the first skin and the second skin together.

[0007] Another embodiment is a system including a circumferential joint. The circumferential joint includes: a splice fitting that straddles a lap join between a first skin of a first circumferential section of the aircraft and a second skin of a second circumferential section of the aircraft; a filler disposed at the second circumferential section that contacts a first portion of the base of the splice fitting; and a flange of a stringer at the first circumferential section that contacts a second portion of the base of the splice fitting, the flange and the filler together defining a flat plane for receiving the splice fitting.

[0008] Other illustrative embodiments (e.g., methods and computer-readable media related to the foregoing embodiments) may be described below. The features, functions, and advantages that have been discussed may be implemented independently in different embodiments or may be combined in other embodiments, and further details of these features, functions, and advantages may be understood with reference to the following description and the drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0009] Some embodiments of the present disclosure will now be described by way of example only and with reference to the drawings. In all the drawings, the same reference numerals denote the same elements or elements of the same type.

[0010] Figure 1 An aircraft in an illustrative embodiment is illustrated.

[0011] Figure 2 Depicts a cross-section of a fuselage section of the Figure 1 fuselage of an aircraft in an illustrative embodiment.

[0012] Figure 3 Is an internal view of a circumferential joint between circumferential sections of a fuselage in an illustrative embodiment.

[0013] Figure 4 Is a side view of a circumferential joint in an illustrative embodiment.

[0014] Figure 5 Is in an illustrative embodiment Figure 8 An enlarged view of a portion of a circumferential joint.

[0015] Figure 6 Is a perspective view of a longitudinal beam that is part of a circumferential joint in an exemplary embodiment.

[0016] Figure 7 Is in an exemplary embodiment Figure 3 Another view of the circumferential joint.

[0017] Figure 8 Is in an exemplary embodiment Figure 3 An enlarged view of the circumferential joint.

[0018] Figure 9 Is an enlarged view of a joint fitting that forms part of a circumferential joint in an exemplary embodiment.

[0019] Figure 10 Is a perspective view of the joint fitting in an exemplary embodiment.

[0020] Figures 11 to 15 Depicts a ply sequence for a skin in an exemplary embodiment.

[0021] Figure 16 Is a flowchart of a method for forming a circumferential joint as illustrated in an exemplary embodiment.

[0022] Figure 17 Is a block diagram of a circumferential joint in an exemplary embodiment.

[0023] Figure 18 Is a flowchart of an aircraft production and maintenance method in an exemplary embodiment.

[0024] Figure 19 Is a block diagram of an aircraft in an exemplary embodiment. Detailed Description

[0025] The accompanying drawings and the following description provide specific exemplary embodiments of the present disclosure. Thus, it should be appreciated that although not explicitly described or shown herein, those skilled in the art will be able to design various arrangements that embody the principles of the present disclosure and are included within the scope of the present disclosure. Moreover, any examples described herein are intended to assist in understanding the principles of the present disclosure and are to be construed as not limited to such specifically recited examples and conditions. As a result, the present disclosure is not limited to the specific embodiments or examples described below, but is limited by the claims and their equivalents.

[0026] Multiple or all of the components discussed herein can be implemented as composite parts. Composite parts (such as carbon fiber reinforced polymer (CFRP) parts) are initially laid-up in multiple layers collectively referred to as a preform. Individual fibers within each layer of the preform can be aligned parallel to each other, but different layers exhibit different fiber orientations in order to increase the strength of the resulting composite part along different dimensions. Also, some layers can include woven fabrics made of fibers. The preform includes a viscous resin that solidifies to harden the preform into a composite part (e.g., for use in an aircraft). Carbon fibers that have been impregnated with uncured thermosetting resin or thermoplastic resin are referred to as "prepreg". Other types of carbon fibers include "dry fibers" that have not been impregnated with thermosetting resin but may include tackifiers or adhesives. Dry fibers are batched with resin prior to hardening. For thermosetting resins, hardening is a unidirectional process referred to as curing, while for thermoplastic resins, if the resin is reheated, it reaches a viscous form, after which the resin can be consolidated into a desired shape and solidified. As used herein, the encompassing term for the process of transforming a preform to its final hardened shape (i.e., transforming the preform into a composite part) is referred to as "hardening", and this term encompasses the curing of thermosetting preforms and the forming / solidifying of thermoplastic preforms into their final desired shape.

[0027] Now, turning to Figure 1 , an illustrative diagram of an aircraft 10 in which the manufacturing systems and methods described herein can be implemented is depicted. In this illustrative example, the aircraft 10 includes wings 15 and 16 that are attached to a fuselage 28 having a nose 12. The aircraft 10 includes an engine 13 attached to wing 15 and an engine 14 attached to wing 16. A tail section 18 is also attached to the fuselage 28. A horizontal stabilizer 20, a horizontal stabilizer 21, and a vertical stabilizer 22 are attached to the tail section 18 of the fuselage 28. The fuselage 28 itself is formed by multiple barrel sections 29-1, 29-2, and 29-3 (commonly referred to as "barrel sections 29"). The barrel sections 29 have been joined together and define the circumference 30 of the fuselage 28. In this embodiment, three different instances of the barrel sections 29 are labeled, but any suitable number of barrel sections 29 can be utilized to form the fuselage 28 as a design choice.

[0028] Figure 2 Depicted in an illustrative embodiment is Figure 1Cross-section of a first circumferential segment 200 (e.g., a barrel-shaped segment 29, a segment forming a ninety-degree arc, or other curved segment) of the fuselage 28 of the aircraft 10. The first circumferential segment 200 is referred to as "circumferential" because it forms part of the perimeter 30 of the fuselage 28. Thus, the first circumferential segment 200 can include Figure 1 One of the barrel-shaped segments 29 of the barrel-shaped segment 29.

[0029] As Figure 2 Shown, the first circumferential segment 200 includes a first skin 210, and stringers 220 and frames 230 attached to the skin. In this embodiment, the first circumferential segment 200 is circular with a center 240.

[0030] Figure 3 Is an internal view of a circumferential joint between circumferential segments of the fuselage in an exemplary embodiment, and corresponds to Figure 2 View arrow 3. The circumferential joint 390 extends around the entire first circumferential segment 200 (e.g., barrel-shaped segment 29-2), however, only a portion of the circumferential joint 390 is shown in this view. As Figure 3 Shown, the circumferential joint 390 is between the first circumferential segment 200 (including the first skin 210, stringers 220, and frames 230) and the second circumferential segment 300 (including the second skin 310, stringers 320, and frames 330). At the circumferential joint 390, the first circumferential segment 200 and the second circumferential segment 300 (e.g., barrel-shaped segment 29-3) are joined together via a lap join 350. The first circumferential segment 200 includes the first skin 210 and stringers 220, while the second circumferential segment 300 includes the second skin 310 and stringers 320. The lap join 350 includes various components that secure the first circumferential segment 200 to the second circumferential segment 300, as discussed below with reference to the drawings. According to embodiments, these different components can be secured via fastening, co-curing, or co-bonding.

[0031] Figure 4 Is a side view of the circumferential joint 390 in an exemplary embodiment, and corresponds to Figure 3 View arrow 4. As Figure 4 Shown, the circumferential joint 390 includes many interconnected components. Generally from Figure 4starting from the bottom and moving upward (i.e., traveling from the outside to the inside), the circumferential joint 390 includes a second skin 310 that forms an interlocking joint 350 (e.g., a ship lap joint, a rabbet join, etc.) with the first skin 210. Specifically, the second skin 310 is disposed below the lip 414 of the first skin 210 and is indexed relative to the shoulder 416 of the first skin 210 via a Determinant Assembly (DA) hole 498, leaving a gap 417 for airtight sealing. In one embodiment, the DA hole 498 is implemented as a pilot hole that aligns the tops, left sides, and right sides of the first skin 210 and the second skin 310. When the pilot hole is aligned with the skins, it is known that the skins are in the desired alignment state.

[0032] The lip 414 and the shoulder 416 are provided at the distal portion 420 of the first skin 210. Closest to or at the circumferential joint 390, the thickness (TB) of the second skin 310 increases at a ramp 424, and the thickness (TA) of the first skin 210 increases at a ramp 412. This increases the strength of the interlocking joint 350. In this embodiment, after the ramps have fully increased the thicknesses of the first skin 210 and the second skin 310, TB is less than TA. This allows sufficient material to remain for the lip 414 after the first skin 210 has been machined by a cutting machine.

[0033] A longitudinal beam 320 having a flange 322 and a body 324 is disposed on top of the second skin 310, while a longitudinal beam 220 having a flange 222 and a body 224 is disposed on top of the first skin 210. The flange 222 extends along the lip 414 toward the second skin 310. A filler 430 is disposed on top of the flange 322 of the longitudinal beam 320. A joint fitting 440 (e.g., an H-shaped fitting, an L-shaped bracket, a T-shaped bracket, or others) partially rests on top of the filler 430 and partially on top of the flange 222. The joint fitting 440 is fixed via a fastener 450, and a low boy 460 and a shear tie joint 470 project upward from the joint fitting 440. Specifically, at position 492, the fastener 450 is driven from the outside to the inside through the second skin 310, the flange 322, the filler 430, and the joint fitting 440. At position 494, the fastener 450 is driven from the outside to the inside through the second skin 310, an optional gasket (e.g., Figure 5 gasket 510), the lip 414, the flange 222, and the joint fitting 440. At position 496, the fastener 450 is driven from the outside to the inside through the first skin 210, the flange 222, and the joint fitting 440.

[0034] The frame 480 (e.g., a composite or titanium frame) has been added to the figure and is secured to the shear tie joint 470. Although the shear tie joint 470 is illustrated as centered above the lap joint 350, in other embodiments, the shear tie joint 470 is positioned to the left or right of the lap joint 350. In still other embodiments, the stringer 320 or the stringer 220 may extend forward or backward beyond the lap joint 350.

[0035] With Figure 4 this arrangement, no joint doubler is required. Thus, a technical benefit is provided in that the labor associated with aligning the joint doubler is saved and the weight at the circumferential joint 390 can be reduced as a result of using fewer components.

[0036] Referring Figure 5 to Figure 5 which provides further details of the circumferential joint 390, which Figure 4 is an enlarged view of the lap joint 350 of Figure 5 in the exemplary embodiment. Specifically, Figure 4 corresponds to region 5 of Figure 5 . In

[0037] this embodiment, an optional spacer 510 has been included between the lip 414 and the second skin 310. The combined thickness of the flange 322 and the filler 430 is equal to the combined thickness of the spacer 510, the lip 414, and the flange 222. This creates a flat plane 530 for receiving the joint fitting 440. The spacer 510 is not a joint doubler because it is not a structural component of the circumferential joint 390 and also because the spacer 510 does not form a single strap joint across two abutting skin segments.

[0038] Figure 6 is a perspective view of the stringers 220 and 320 covering a portion 600 of the circumferential joint 390 in the exemplary embodiment and corresponds to view arrow 6 of Figure 4 . In Figure 6 , the flange 222 of the stringer 220 extends beyond the body 224 of the stringer 220 and along the lip 414 across the lap joint 350. At this time, the flange 322 of the stringer 320 is covered by the filler 430. The top surface 622 of the flange 222 and the top surface 632 of the filler 430 are coplanar with each other and define Figure 5a flat plane 530 such that the joint fitting 440 can straddle these components during installation. The orientation of the plane 530 along the perimeter 30 of the fuselage 28 changes as the perimeter 30 is traversed to accommodate the local geometry of the perimeter 30 (i.e., the curvature of the perimeter 30).

[0039] Figure 7 is in the exemplary embodiment Figure 3 of the circumferential joint 390 and corresponds to Figure 3 view arrow 7. Specifically, Figure 7 a different perspective of the circumferential joint 390 is provided and additionally depicts the frame 480 secured to the circumferential joint 390.

[0040] Figures 8 to 10 is an additional view providing additional background on the components stated above. Figure 8 is in the exemplary embodiment Figure 3 of the enlarged circumferential joint 390 and corresponds to Figure 7 view arrow 8. Specifically, Figure 8 background is provided by depicting the joint fitting 440, the stub 460, and the shear tie joint 470 from a new perspective. In this view, the frame 480 is omitted. Figure 9 is an enlarged view of the joint fitting 440 forming part of the circumferential joint 390 in the exemplary embodiment and corresponds to Figure 8 region 9. In this view, the frame 480 is included and the frame 480 overlaps the joint fitting 440, the stub 460, and the shear tie joint 470.

[0041] Figure 10 is a perspective view of the joint fitting 440 in the exemplary embodiment. Figure 10 Shows that the joint fitting 440 is symmetric about line 1010 but asymmetric about line 1020 (i.e., front to back). Specifically, a first portion 1030 of the base 442 (e.g., the lower side) of the joint fitting 440 rests on the filler 430 and has a different shape from a second portion 1040 of the base 442 that rests on or otherwise contacts the flange 222 of the stringer 220. This front-to-back shape difference accommodates placing the joint fitting on the filler 430 (e.g., on the rear side) and on the flange 222 of the stringer 220 (e.g., on the front side). It should be understood that the completed circumferential joint 390 will include joint fittings 440 along the entire perimeter 30 of the fuselage 28 to straddle the interlocking joint 350 and enhance strength.

[0042] Utilizing with respect to Figures 1 to 10The discussion of the overall shape of the circumferential joint 390 provided focuses further discussion on the ply-by-ply stacking of the first skin 210 and the method for manufacturing the circumferential joint 390 discussed herein.

[0043] Figure 11 Illustrated is a ply sequence 1100 of the first skin 210 in an exemplary embodiment. The thickness of the first skin 210 increases from left to right as additional plies are incorporated. Specifically, the sequence 1100 indicates the number and fiber orientation of plies 1130 of a fiber reinforcement 1140 (e.g., CFRP), which is made of unidirectional fiber tows set at a desired fiber angle. In this embodiment, the angles include 90 degrees (as shown by fiber orientation 1132), 0 degrees (as shown by fiber orientation 1134), +45 degrees (as shown by fiber orientation 1136), and -45 degrees (as shown by fiber orientation 1132). Design constraints may impose certain requirements on the sequence 1100. For example, the design constraints may require that the fiber orientation be symmetric along the entire thickness 1122 of the sequence 1100 (e.g., with respect to the centerline 1120). This presents a difficulty because the first skin 210 will have a lip 414 machined therefrom, and the lip will have a different thickness than the remainder of the first skin 210.

[0044] To address this difficulty, the sequence 1100 exhibits "double symmetry", where the region 1112 corresponding to the lip 414 exhibits symmetry in fiber orientation about its centerline 1110, while the entire thickness 1122 of the ply sequence 1100 corresponding to the combined thickness of the lip 414 and the shoulder 416 also exhibits symmetry, but with respect to the centerline 1120. This enables the first skin 210 to meet the design requirements regarding symmetry of fiber orientation across the centerline even when the lip 414 is machined from the first skin 210.

[0045] Figures 12 to 15 Illustrated are additional sequences of a "double symmetric" arrangement of plies. In Figures 12 to 15 a uniform ply length is provided for each ply. However, a ramp technique similar to the ramp technique provided in Figure 11 may be utilized to increase the thickness of the first skin 210 from left to right. In Figure 12In [sequence 1200], laminae 1130 with fiber orientations 1132, 1134, 1136, and 1138 are provided to form a region 1212 corresponding to the lip 414. Region 1212 is a portion remaining after machining away a part (e.g., one-third) of the thickness 1222 of sequence 1200. Region 1212 is symmetric about the centerline 1210, and the thickness 1222 of sequence 1200 is symmetric about the centerline 1220.

[0046] In Figure 13 [sequence 1300], an additional lamina 1350 of a woven fabric has been integrated. Sequence 1200 provides laminae 1130 with fiber orientations 1132, 1134, 1136, 1138, and 1140 to form a region 1312 corresponding to the lip 414. Region 1312 is a portion remaining after machining away a part (e.g., half, one-third, two-thirds, or any fraction) of the thickness 1322 of sequence 1300. Region 1312 is symmetric about the centerline 1310, and the thickness 1322 of sequence 1300 is symmetric about the centerline 1320.

[0047] In Figure 14 [sequence 1400], a sequence 1130 of laminae with the following fiber orientations is provided: 0 degrees (e.g., fiber orientation 1432), 60 degrees (e.g., fiber orientation 1434), and -60 degrees (e.g., fiber orientation 1436). Sequence 1400 exhibits double symmetry. Specifically, region 1412 is symmetric about the centerline 1410, and the thickness 1422 of sequence 1400 is symmetric about the centerline 1420.

[0048] In Figure 15 [sequence 1500], another sequence 1130 of laminae with the following fiber orientations is provided: 0 degrees (e.g., fiber orientation 1432), 60 degrees (e.g., fiber orientation 1434), and -60 degrees (e.g., fiber orientation 1436). Sequence 1500 exhibits double symmetry. Specifically, region 1512 is symmetric about the centerline 1510, and the thickness 1522 of sequence 1500 is symmetric about the centerline 1520. Exemplary details of forming the circumferential joint 390 will be Figure 16 discussed. For this embodiment, it is assumed that the first circumferential segment 200 and the second circumferential segment 300 are waiting to be joined together.

[0049] Figure 16FIG. is a flowchart of a method 1600 for forming a circumferential joint 390 illustrated in an exemplary embodiment. The steps of method 1600 are described with reference to the components discussed in the figures above, but those skilled in the art will recognize that method 1600 may be performed in other systems. The steps of the flowchart described herein do not include all steps and may include other steps not shown. The steps described herein may also be performed in an alternative order.

[0050] In step 1602, a first skin 210 of a first circumferential segment 200 of the fuselage 28 is formed. The first skin 210 includes a distal portion 420 that includes a lip 414 and a shoulder 416. In optional step 1604, a portion of the thickness of the first skin 210 is machined to form the lip 414 and the shoulder 416. This operation can be performed by a mill, a cutting machine, or other suitable machinery (such as Figure 17 a machining tool 1790).

[0051] Optionally, step 1606 includes inserting a spacer 510 between the lip 414 of the second circumferential segment 300 of the fuselage 28 and the second skin 310.

[0052] Step 1608 includes aligning the second skin 310 of the second circumferential segment 300 with the shoulder 416 such that the lip 414 overlaps the second skin 310. This operation is performed circumferentially along the entire second circumferential segment 300. The result is an interlocking joint 350 (e.g., a lapped joint, a notched joint, etc.) where the second skin 310 is seated against the lip 414 and the shoulder 416 of the first skin 210. Thus, although the second skin 310 is seated against the first skin 210 or otherwise aligned with the first skin, the final result is not a butt joint.

[0053] Step 1610 includes securing the first skin 210 and the second skin 310 together via the circumferential joint 390.

[0054] In one embodiment, step 1610 includes aligning a filler 430, a joint fitting 440, a stub 460, and / or a shear tie joint 470; and securing these components via fasteners, co-curing, or co-bonding. Optionally, step 1612 includes placing the filler 430 onto a flange 322 of the stringer 320 that is attached to the second skin 310.

[0055] Optionally, step 1614 includes: cross - mounting the joint fitting 440 across the first skin 210 and the second skin 310. In one embodiment, this includes: placing a second portion 1040 of the base 442 of the joint fitting 440 onto the flange 222 of the stringer 220 at the first skin 210; and then sliding the filler 430 between a first portion 1030 of the base 442 of the joint fitting 440 and the flange 322 of the stringer 320.

[0056] Optionally, step 1616 includes: aligning the filler 430 with an end 520 of the lip 414 such that the combined thickness of the filler 430 and the flange 322 of the stringer 320 at the second circumferential segment 300 of the fuselage 28 is equal to the combined thickness of the lip 414, the gasket 510 contacting the lip 414, and the flange 222 of the stringer 220 at the first circumferential segment 200. In one embodiment, this includes: sliding the filler 430 under the joint fitting 440 and on top of the flange 322 such that the filler 430 is clamped between the joint fitting 440 and the flange 322.

[0057] Optionally, step 1618 includes: driving a fastener 450 through the first skin 210, the flange 222, and / or the joint fitting 440 at position 496. Optionally, step 1620 includes: driving a fastener 450 through the second skin 310, the flange 430, the filler 430, and / or the joint fitting 440 at position 492.

[0058] Upon completion of the circumferential joint 390, additional structures can be attached. For example, optionally, step 1622 includes: installing a frame 480 at the circumferential joint 390. In some other embodiments, the frame 480 is installed as part of step 1610.

[0059] Optionally, step 1624 includes: forming the circumferential joint 390 without using a joint doubler (i.e., foregoing / omitting the installation of the joint doubler). Although not illustrated, a joint doubler is a structural component of a joint that overlaps two abutting (or otherwise aligned) skins to form a single - doubler butt joint. The circumferential joint 390 discussed herein obviates the need for a joint doubler because the lip 414 serves the strengthening structural purpose that would otherwise be performed by the joint doubler.

[0060] Relative to existing systems that use joint doublers to form a single - doubler butt joint, method 1600 provides a technical benefit by providing a strengthened joint between segments of the fuselage 28. Moreover, method 1600 reduces the amount of labor involved in aligning the components of the circumferential joint 390 (because alignment with a joint doubler is not required) and reduces weight (because the joint doubler is eliminated). This results in benefits related to assembly and reduced fuel - consumption costs.

[0061] Example

[0062] In the following examples, additional processes, systems, and methods are described in the context of a circumferential joint of a circumferential segment of a fuselage in an exemplary embodiment.

[0063] Figure 17 is a block diagram of a circumferential joint 1700 between a first circumferential segment 1750 and a circumferential segment 1710 of a fuselage 28 in an exemplary embodiment. In this embodiment, the first circumferential segment 1750 includes a first skin 1760 having a lip 1762 and a shoulder 1764, and also includes a stringer 1752 having a flange 1754. The second circumferential segment 1710 includes a second skin 1720 that is aligned with the shoulder 1764 (e.g., butted against or nested at), and also includes a stringer 1712 having a flange 1714. A filler 1730 rests on top of the flange 1714, and a joint fitting 1770 rests on top of the filler 1730 and the flange 1754 of the stringer 1752. Fasteners 1792 hold the joint fitting 1770 in place. A stub 1782 is attached to the joint fitting 1770, and so is a shear tie joint 1784. A frame 1780 is attached to the shear tie joint 1784. Moreover, a gasket 1702 is disposed between the lip 1762 and the second skin 1720. Figure 17 A machining tool 1790 for machining the lip 1762 from the thickness (T3) of the first skin 1760 is also depicted. The machining tool 1790 can include, for example, a milling machine, a circular saw, or a reciprocating saw or other tools.

[0064] Figure 17 It is also illustrated that the combined thickness (T2) of the filler 1730 and the flange 1714 of the stringer 1712 at the second circumferential segment 1710 of the fuselage 28 is equal to the combined thickness (T1) of the lip 1762, the gasket 1702 contacting the lip 1762, and the flange 1754 of the stringer 1752 at the first circumferential segment 1750. Moreover, the combined thickness (T4) of the second skin 1720 and the gasket 1702 is equal to the thickness (T5) of the shoulder 1764.

[0065] More particularly with reference to the drawings, embodiments of the present disclosure can be in a method 1800 of aircraft manufacturing and maintenance as shown in Figure 18 and as shown in Figure 19described in the context of the aircraft 1802 shown. During pre-production, method 1800 may include the specification and design 1804 of the aircraft 1802 and the procurement of materials 1806. During production, the fabrication 1808 of components and sub-components of the aircraft 1802 and the system integration 1810 may be performed. Thereafter, the aircraft 1802 may undergo certification and delivery 1812 for use 1814. When in use by a customer, routine operational maintenance and servicing 1816 (which may also include modification, reconfiguration, refurbishment, etc.) may be scheduled for the aircraft 1802. The apparatus and methods specifically implemented herein may be used during any one or more suitable stages of production and servicing described in method 1800 (e.g., specification and design 1804, procurement of materials 1806, fabrication 1808 of components and sub-components, system integration 1810, certification and delivery 1812, in use 1814, maintenance and servicing 1816) and / or any suitable components of the aircraft 1802 (e.g., airframe 1818, systems 1820, interior 1822, propulsion system 1824, electrical system 1826, hydraulic system 1828, environmental system 1830).

[0066] Each process in the processes of method 1800 may be performed or implemented by a system integrator, a third party, and / or an operator (e.g., a customer). For the purposes of this description, a system integrator may include, but is not limited to, any number of aircraft manufacturers and prime system subcontractors; a third party may include, but is not limited to, any number of vendors, subcontractors, and suppliers; and an operator may be an airline, a leasing company, a military entity, a service organization, etc.

[0067] As Figure 19 shown, an aircraft 1802 produced in accordance with method 1800 may include an airframe 1818 having a plurality of systems 1820 and an interior 1822. Examples of systems 1820 include one or more of the following: a propulsion system 1824, an electrical system 1826, a hydraulic system 1828, and an environmental system 1830. Any number of other systems may be included. Although an aerospace example is shown, the principles of the present invention may be applied to other industries such as the automotive industry.

[0068] As previously mentioned, the devices and methods specifically implemented herein can be used during any one or more suitable stages of the production and maintenance described in method 1800. For example, components or sub-components corresponding to component and sub-component manufacturing 1808 can be fabricated or manufactured in a manner similar to that of components or sub-components produced while the aircraft 1802 is in use. Moreover, during sub-component manufacturing 1808 and system integration 1810, one or more device implementations, method implementations, or combinations thereof can be utilized, for example, by significantly accelerating the assembly of the aircraft 1802 or reducing the cost of the aircraft. Similarly, when the aircraft 1802 is in use (e.g., and without limitation, during maintenance and servicing 1816), one or more of the device implementations, method implementations, or combinations thereof can be utilized. Thus, the present invention can be used in any stage or any combination thereof discussed herein, such as specification and design 1804, material procurement 1806, component and sub-component manufacturing 1808, system integration 1810, certification and delivery 1812, in use 1814, maintenance and servicing 1816, and / or any suitable components of the aircraft 1802 (e.g., airframe 1818, systems 1820, interiors 1822, propulsion systems 1824, electrical systems 1826, hydraulic systems 1828, and / or environmental systems 1830).

[0069] In one embodiment, the component comprises a portion of the airframe 1818 and is manufactured during component and sub-component manufacturing 1808. The component can then be assembled into the aircraft during system integration 1810 and then utilized during in use 1814 until wear renders the component unusable. Then, during maintenance and servicing 1816, the component can be discarded and replaced with a newly manufactured component. The components and methods of the invention can be utilized throughout component and sub-component manufacturing 1808 to manufacture new components.

[0070] Although specific embodiments have been described herein, the scope of the present disclosure is not limited to those specific embodiments. The scope of the present disclosure is defined by the appended claims and any equivalents thereof.

Claims

1. A method for assembling an airframe of an aircraft, the method comprising the steps of: Forming a first skin of a first circumferential section of a fuselage, the first circumferential section including the first skin and a first stringer, wherein the first skin includes a distal portion, the distal portion including a lip and a shoulder; Aligning a second skin of a second circumferential section of the fuselage with the shoulder such that the second skin overlaps the lip to form an interlocking joint, wherein the second circumferential section includes the second skin and a second stringer; and Securing the first skin and the second skin together via a circumferential joint by the steps of: Placing a filler onto a flange of the second stringer; Aligning the filler with an end of the lip, wherein a combined thickness of the filler and the flange of the second stringer is equal to a combined thickness of the lip and a flange of the first stringer to form a flat plane; Straddling an interlocking joint between the first stringer and the second stringer on the flat plane with a joint fitting; Driving fasteners through the first skin, the flange of the first stringer, and the joint fitting; and Driving fasteners through the second skin, the flange of the second stringer, the filler, and the joint fitting.

2. The method according to claim 1, wherein: The step of forming the first skin includes: machining away a portion of the thickness of the first skin to form the lip and the shoulder.

3. The method according to claim 1, the method further comprising the step of: Inserting a spacer between the lip and the second skin.

4. The method according to claim 1, the method further comprising the step of: Forming the circumferential joint without using a joint doubler.

5. The method according to claim 1, the method further comprising the step of: Installing a frame at the circumferential joint.

6. A system including a portion of an airframe of an aircraft, the system comprising: A first circumferential section of a fuselage, the first circumferential section including a first skin and a first stringer, wherein the first skin includes a distal portion, the distal portion including a lip and a shoulder; A second circumferential section of the fuselage, the second circumferential section including a second skin and a second stringer; And A circumferential joint that secures the first circumferential section to the second circumferential section, Wherein the circumferential joint includes: An interlocking joint where the second skin overlaps the lip of the first skin; A filler that is placed onto a flange of the second stringer and aligned with an end of the lip such that a combined thickness of the filler and the flange of the second stringer is equal to a combined thickness of the lip and a flange of the first stringer to form a flat plane; A joint fitting that straddles the interlocking joint between the first stringer and the second stringer on the flat plane; Fasteners driven through the first skin, the flange of the first stringer, and the joint fitting; and Fasteners driven through the second skin, the flange of the second stringer, the filler, and the joint fitting.

7. The system according to claim 6, wherein: the fiber orientation of the fiber-reinforced material ply in the first skin is symmetric about the centerline of the first skin; and the fiber orientation of the fiber-reinforced material ply in the lip is symmetric about the centerline of the lip.

8. The system according to claim 6, wherein: the filler abuts against the end of the lip.

9. The system according to claim 6, wherein: the joint fitting is asymmetric from front to back.

10. The system according to claim 6, the system further comprising: a gasket disposed between the lip and the second skin.

11. The system according to claim 6, wherein: the flange of the first longitudinal beam extends along the lip towards the second skin.

12. The system according to claim 6, wherein: the first skin includes a ramp that increases the thickness of the first skin at the circumferential joint; the second skin includes a ramp that increases the thickness of the second skin at the circumferential joint.

13. The system according to claim 6, wherein: the first circumferential segment of the fuselage includes a barrel segment; and the second circumferential segment of the fuselage includes a barrel segment.

14. The system according to claim 6, wherein: the thickness of the gasket and the second skin is equal to the thickness of the shoulder.

15. A system including a circumferential joint, the system comprising: a first circumferential segment of a fuselage, the first circumferential segment including a first skin and a first longitudinal beam, wherein the first skin includes a distal portion that includes a lip and a shoulder; a second circumferential segment of the fuselage, the second circumferential segment including a second skin and a second longitudinal beam; the circumferential joint that fixes the first circumferential segment to the second circumferential segment, the circumferential joint including: an interlocking joint, where the second skin overlaps the lip of the first skin at the interlocking joint; a filler placed on the flange of the second longitudinal beam and aligned with the end of the lip such that the combined thickness of the filler and the flange of the second longitudinal beam is equal to the combined thickness of the lip and the flange of the first longitudinal beam to form a flat plane; a joint fitting that straddles the interlocking joint between the first longitudinal beam and the second longitudinal beam on the flat plane; fasteners driven through the first skin, the flange of the first longitudinal beam, and the joint fitting; and fasteners driven through the second skin, the flange of the second longitudinal beam, the filler, and the joint fitting.

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