• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    circumference splicing to join carcass structures. The present invention relates to a splicing of the housing structure and method which may include a first panel (110a) having a first edge (120a), a second panel having a second edge, said second edge being positioned in alignment towards the edge with said first edge (120a) to form a seam joint, a strip (12) bridging the seam joint and is secured to the first panel (110a) and the second panel, wherein the strip (12) is provided with a first tapered region and a second tapered region, a first socket is provided with a tapered section and a flat section, the tapered section being attached to the first tapered region of the strip (12), and a second socket is provided with a tapered section and a flat section, the tapered section being attached to the second tapered region of the strip (12).
    公开号:BR102013027906B1
    申请号:R102013027906-4
    申请日:2013-10-30
    公开日:2022-01-11
    发明作者:Paul Diep；Bernhard Dopker
    申请人:The Boeing Company；
    IPC主号:
    专利说明:

    FIELD
    [001] The present description generally refers to the joining of shell structures and, more particularly, to a splicing joint for joining adjacent sections of a shell structure, such as the fuselage sections of an aircraft. BACKGROUND
    [002] The primary structural elements of large aircraft are typically made of metal or composite materials. For example, the fuselage shells of such an aircraft can typically be manufactured from high strength aluminum alloy materials or fiber reinforced resin materials which have relatively high strength-to-weight ratios.
    [003] An aircraft may include two or more fuselages, i.e., shell sections, which are point-connected and joined in a circumferential manner at a splice joint to interconnect the fuselage sections and form the complete fuselage structure. To precisely install the integration joints, the fuselage sections are aligned and fastening holes are drilled through the connecting splice plates and the underlying housing structure. For example, adjacent fuselage sections may be joined by a seam consisting of a frame or bulkhead that is positioned within the fuselage and on bridges between the reinforcing spars of the adjacent fuselage section. A plurality of slots extend through the frame and facilitate continuity of fuselage sections. As another example, adjacent fuselage sections may be integrated together by the splicing plate and the plurality of splicing fittings secured between a pair of adjacent fuselage sections to bridge the interface between adjacent fuselage sections. Generally, the splice plate is flat and forms a flat splicing cord that extends across the splice joint. Depending on the type of splice joint, assembly may require a plurality of mouse holes arranged in the frame or bulkhead through which the splice fittings extend.
    [004] Typically, a splice fitting can be formed from a metallic material such as titanium. Since titanium is a relatively expensive material, the material and manufacturing costs associated with a titanium splice fitting can add to the costs associated with producing the aircraft. Additionally, the fuselage sections may be joined by a plurality of fasteners extending through the frame to connect the frame to the fuselage sections and a plurality of fasteners extending through the splice fittings to connect the fittings to the fuselage sections. In order to install the fasteners, a plurality of fastening holes must be drilled through the seam to receive the fasteners. A disadvantage of such an assembly process is that drilling holes through a stack of different materials that contain titanium splice fittings takes a significant amount of time, thereby increasing the time required to assemble the fuselage sections as well as the cost of repair. labor associated with such assembly. Furthermore, holes drilled through a splice fitting made of titanium or other metallic material often require the parts to be pulled apart and deburred, thereby adding to the time and costs associated with assembling the fuselage. A further disadvantage of such an assembly process is that shims may be required to align the fuselage section and fit the splice fittings along the joint interface, thereby adding to even more time and cost.
    [005] In this way, those skilled in the art continue with research and development efforts in the field of joining shell structures, particularly in the field of aircraft assembly. SUMMARY
    [006] In one embodiment, the splice of the disclosed housing structure may include a first panel having a first edge, a strip attached to the first panel and extending beyond the first edge, the strip having a first region tapered, and a first socket having a tapered section and a flat section, the tapered section being attached to the first tapered region of the strip. A frame is optionally connected to the strap.
    [007] In another embodiment, the seam of the disclosed carcass structure may include a first panel having a first edge, a second panel having a second edge, the second edge being positioned in alignment towards the edge with the first edge to form a seam joint, a strip bridging the seam joint and is attached to the first panel and the second panel, the strip being provided with a first tapered region and a second tapered region, a first socket is provided with a tapered section and a flat section, the tapered section being attached to the first tapered region of the strip, and a second socket having a tapered section and a flat section, the tapered section being secured to the second tapered region of the strip.
    [008] In another embodiment, the seam disclosed for joining adjacent panels of a housing structure may include a strip having a first side, a second side, a generally flat intermediate region, a first tapered region that extends downwards. proximate the intermediate region to approximate the first side, and a second tapered region extending downwardly close to the intermediate region to approximate the second side, a first socket provided with a tapered section and a flat section, the tapered section being secured to the first tapered region of the strip, and a second socket which is provided with a tapered section and a flat section, the tapered section being attached to the second tapered region of the strip. A Z-section frame or a C-section can be attached to the middle flat section of the splicing strip.
    [009] In another embodiment, a method of joining a carcass structure is revealed, the method may include the steps of: (1) providing a first panel of the carcass structure provided with a first edge, (2) providing a strip having a first tapered region and a second tapered region, (3) positioning the strip adjacent the first edge such that the first tapered region is adjacent to the first panel and the second tapered region extends away from the first edge, (4) securing the strip to the first panel, (5) providing a first groove having a tapered section and a flat section, (6) positioning the first groove such that the tapered section is adjacent to the first tapered region of the strip and the flat section is adjacent to the first panel, (7) attaching the first groove to the strip, and (8) attaching the first groove to the first panel.
    [0010] In yet another embodiment, a method of joining a carcass structure is disclosed, the method may include the steps of: (1) providing a first panel of the carcass structure having a first edge, (2) providing a strip having a first tapered region and a second tapered region, (3) positioning the strip adjacent the first edge such that the first tapered region is adjacent to the first panel and the second tapered region extends away from the first edge, (4) securing the strip to the first panel, (5) providing a first groove having a tapered section and a flat section, (6) positioning the first groove such that the tapered section is adjacent to the first tapered region of the strip and the flat section is adjacent to the first panel, (7) attaching the first recess to the strip, (8) attaching the first recess to the first panel, (9) providing a second panel of the housing structure provided with a second edge, (10) positioning the second panel adjacent to the first panel l such that the first and second edges are in alignment towards the edge forming a seam joint and the second tapered section is adjacent to the second panel and the strip bridges the seam joint, (11) attach the strip to the second panel , (12) providing a second groove having a tapered section and a flat section, (13) positioning the second groove such that the tapered section is adjacent to the second tapered region of the strip and the flat section is adjacent to the second panel, (14) ) attaching the second notch to the strip, and (15) attaching the second notch to the second panel.
    [0011] Other embodiments of the revealed grafted bonding surface will become apparent from the detailed description, the accompanying drawings and the following appended claims. BRIEF DESCRIPTION OF THE DRAWINGS
    [0012] Figure 1 is a perspective view of an aircraft assembled in accordance with the present description;
    [0013] Figure 2 is an exploded perspective view of a plurality of fuselage sections of the aircraft of Figure 1;
    [0014] Figure 3 is a partial perspective view of a seam between adjacent fuselage sections of the aircraft of Figure 1;
    [0015] Figure 4 is a partial perspective view of the single seam of the seam joint of Figure 3;
    [0016] Figure 5 is a partially exploded side perspective view of the seam of Figure 4;
    [0017] Figure 6 is a partially exploded top perspective view of the seam of Figure 4;
    [0018] Figure 7 is a top view of the splice of Figure 4;
    [0019] Figure 8 is a side view of the seam of Figure 4;
    [0020] Figure 9 is a partial exploded perspective view of the splice; and
    [0021] Figure 10 is a cross-sectional view of a splicing plate of the revealed splice. DETAILED DESCRIPTION
    [0022] The following detailed description refers to the accompanying drawings, which illustrate specific embodiments of the description. Other modalities endowed with different structures and operations do not depart from the scope of the present description. Equal numerical references may refer to the same element or component in different drawings.
    [0023] Referring to Figures 1 and 2, an aircraft 100 may include a fuselage 102 provided with a plurality of fuselage sections 104 (individually identified as fuselage sections 104a-e). The fuselage sections 104 may be joined by a plurality of corresponding splice joints 106 (individually identified as splice joints 106a-e).
    [0024] Each fuselage section 104 may include a composite skin 108 that extends three hundred and sixty degrees (360°) around a longitudinal axis A of the fuselage 102. Throughout this description, the term "fuselage section" is used for convenience to refer to any housing structure that extends three hundred and sixty degrees (360°) around an axis. One skilled in the art will appreciate that a fuselage section 104 may not be limited to generally cylindrical structures, but may include structures that are circular, elliptical, oval, oval, straight, tapered, or other shapes in cross section. Furthermore, one skilled in the art can appreciate that the fuselage sections 104 may be one-piece sections wherein the skins 108 are one-piece skins that extend continuously three hundred and sixty degrees (360°) around the axis or axis. may be formed from two or more skin segments joined together to form the complete three hundred and sixty degree (360°) fuselage section 104.
    [0025] The fuselage 102 may additionally include a passenger cabin configured to hold a plurality of passenger seats. Each fuselage section 104 may also include a plurality of window cutouts (not shown) to provide passengers seated in the passenger cabin with views out of the aircraft 100. The fuselage 102 may also include passenger doors, cargo doors, antennas and similar.
    [0026] Referring to Figure 3, a seam, generally designated 10, may bridge between and structurally join a pair of adjacently positioned fuselage sections 104 defining seam joint 106 therebetween. This view faces outward on a portion of a splice joint 106 from within the fuselage 102. The splices 10 may generally be located at opposite ends of the fuselage sections 104. The splice 10 may be positioned within the fuselage sections 104 to integrate the fuselage sections 104 and provide strength and stability to the resulting fuselage 102. Although only a portion of the circumferentially extending seam 10 is illustrated in Figure 3, the seam may extend around the entire inner circumference of the fuselage sections 104 or may extend around only one or more parts thereof.
    [0027] Each fuselage section 104 may include a panel 110 (individually identified as a first panel 110a and a second panel 110b). A first fuselage section 104a may include a first panel 110a and a second fuselage section 104b may include a second panel 110b positioned in edge alignment with the first panel 110a. In one embodiment, the panels 110 may be at least generally similar in structure and function to meet the panel portions of the aircraft's fuselage sections. For example, panel 110 may include a plurality of stiffener or stiffener spars 114 attached to skin 108. Skin 108 may be secured and covers the plurality of stiffener spars 114 to form the outer surface of fuselage section 104.
    [0028] Referring to Figures 4 to 8, each reinforcing spar 114 may include a raised portion 116 that projects away from the cladding 108 and a plurality of flange parts 118 attached directly to the cladding 108. In the illustrated embodiment, stiffener stringers 114 may have generally "U" shaped cross sections. However, one skilled in the art will appreciate that the gussets 114 may have other shapes in cross section, including "L" shapes, "C" shapes, inverted "T" shapes, "I" shapes, etc. . In still other embodiments, panels 110 may include other features; as support elements or offsets compatible with the flange parts 118 of the stiffener spars 114.
    [0029] Reinforcement stringers 114 may be positioned on skinning 108 so that the flange portion 118 of a brace stringer 114 is aligned with a corresponding flange portion 118 of an adjacent brace member 114. By aligning the flange parts 118 in such a manner, the flange portions 118 may form a plurality of at least approximately continuous support surfaces that extend between the raised portions 116 of the gussets 114. In another embodiment, as illustrated, the gussets 114 may be spaced from each other in the skin 108 so that the flange portion 118 of a gusset 114 is circumferentially offset (or otherwise spaced) from a corresponding flange portion 118 of an adjacent gusset 114.
    [0030] Referring next to Figures 4 to 8, which depict a segment of the splice joint 106 to more clearly illustrate the disclosed splice 10. The splice 10 may include a strip 12 that extends at least partially circumferential around the splice joint 106 of the fuselage sections 104. The strip 12 may have a width, as defined in the longitudinal direction of the fuselage 102, that is sufficient to bridge across the parts and cover the parts of both fuselage sections. adjacent fuselage 104.
    [0031] Strip 12 may include a bottom surface 14 to make contact with skin 108 of adjacent panels 110. Bottom surface 14 may be curved and the curvature of the bottom surface may correspond to the curvature of fuselage 102 (Figure 2). ). A flat bottom surface 14 is also contemplated.
    [0032] Opposite sides 16 (identified individually in Figures 4 and 5 as first side 16a and second side 16b) of strip 12 may extend away from splice joint 106 and can be secured to an edge 120 (individually identified in Figures 4 and 5 as first edge 120a and second edge 120b) of skin 108. A top surface 18 of seam strip 12 may include a generally flat intermediate region 20 and opposite downwardly tapered regions 22 (identified individually in Figures 4 to 6). as first tapered region 22a and second tapered region 22b) that extend near (i.e., at or near) intermediate region 20 to near sides 16. As illustrated, in one embodiment, splicing strip 12 may have a generally isosceles trapezoid cross-sectional shape; however, one skilled in the art will appreciate that the strip 12 may have other generally trapezoidal or other polygonal shapes.
    [0033] As will be described in more detail herein, in one embodiment, the strip 12 may be formed of a composite material, such as a carbon fiber reinforced polymer (CFRP), a composite fiber reinforced polymer, a graphite- epoxy, or similar material, to provide the strength and stability required to structurally join adjacent fuselage sections 104 while reducing material costs, weight, drilling and assembly time over more conventional splices that use metallic splicing parts. In other embodiments, the strip 12 may include other materials, including metallic materials such as aluminum, titanium, steel, etc.
    [0034] Strip 12 may be secured near first edge 120a of skin 108 of first panel 110a and close to second edge 120b of skin 108 of second panel 110b to splice panels 110. Strip 12 may extend continuously or at least approximately continuously around splice joint 106. Alternatively, strip 12 may be segmented around splice joint 106; for example, splice joint 106 may include a plurality of segments of strip 12. Strip 12 may be attached to an inner side of respective skins 108 to maintain a smooth, aerodynamic surface on the exterior of fuselage 102. Strip 12 may be secured to the skins 108 by a plurality of fasteners that extend through the strip 12 and the skins 108. Alternatively, the strip 12 may be attached to the skins 108 or bonded and secured to the skins 108.
    [0035] The strip 12 may be at least approximately as thick as the skins 108, but thicker than the adjacent flange portions 118 of the gussets 114. In one embodiment, as illustrated, the gussets 114 are not extend completely to the edge 120, the skin 108 and sides 16 of the strip 12 may extend near the terminal end of the gussets 114. In other embodiments, the gussets 114 may extend to the edge 120 of the skin 108 and finish it. If necessary, to avoid a step between adjacent surfaces, shim pads or infills (not shown) may be positioned on portions of flange 118 adjacent to strip 12. Infills may include composite materials, metallic materials, or similar materials. In other embodiments, the strip 12, liners 108 and flange portions 118 may have other relative thicknesses such that padding may not be necessary.
    [0036] The splice 10 may also include a plurality of splice sockets 24 (individually identified as a first socket 24a and a second socket 24b) which are disposed on the strip 12 and which extend in a longitudinal direction away from the splice joint. splice 106. Each socket 24 may include a first end 26, a second end 28 and longitudinal sides 30. A base 32 of the socket 24 may include an upwardly tapered section 34 extending near the first end 26 and a generally flat section. 36 which extends near the tapered section 34 to near the second end 28. The sockets 24 may be formed of a composite material such as a carbon fiber reinforced plastic or similar material such as the same composite that forms the strip 12. Alternatively , the sockets 24 may be formed from a metal or metal alloy.
    [0037] In the illustrated embodiment, the upwardly tapered section 34 of the socket 24 may include an upward (i.e. opposite) complementary angle correspondingly with respect to the angle of the downwardly tapered region 22 of the strip 12, such as that the tapered section 34 of the bottom surface 32 can make flush contact with the tapered region 22 of the top surface 18 of the strip. The tapered section 34 may have a width, as defined in the longitudinal direction of the fuselage 102, which is sufficient to extend through the tapered region 22 of the strip 12 and cover the region. Flat section 36 of groove 24 may extend across side 16 of strip 12 to cover skin 108 of panel 110. One groove 24 may be attached to both tapered regions 22 of strip 12 such that a pair of grooves 24 extends longitudinally beyond opposing sides 16 of strip 12 away from seam 106 so as to cover adjacent fuselage sections 104.
    [0038] As shown in the illustrated embodiment, each socket 24 may include a first longitudinal element 38 and a second longitudinal element 40 that extend beyond the sides 16 of the strip 12. Generally, the longitudinal elements 38, 40 may define the flat section 36 of the socket 24. The longitudinally extending members 38, 40 may be configured to cover and receive at least a portion of the reinforcing stringer 114 of a respective panel 110.
    [0039] For example, in one embodiment, each socket 24 may be generally "U" shaped such that each longitudinal member 38, 40 may be spaced from the other and may cover a respective reinforcing stringer 114. The sockets 24 of this embodiment may include the tapered section 34 that extends across the width of the tapered region 22 of the strip 12 and the longitudinal elements 38, 40 that extend out of the tapered section 34 beyond the side 16 of the strip 12. However, the tapered section 34 of the socket 24 may have other widths with respect to the tapered region 22 of the strip 12 and as such, the tapered section 22 of the socket 12 may not need to extend across the entire width of the tapered region 22 of the strip 12. Alternatively, the tapered section 34 of the socket 24 may extend beyond the tapered region 22 of the strip 12 in other embodiments.
    [0040] In other embodiments, the flat section 36 of the socket 24 may include a single longitudinal element that extends beyond the side 16 of the strip 12. In such an embodiment, the flat section 36 may be adequately sized, i.e., the distance between the longitudinal sides 30, to fit between the raised portions 116 of the adjacent reinforcement spars 114.
    [0041] The splice 10 may include a plurality of grooves 24 which overlap different circumferentially spaced portions of the strip 12. The plurality of grooves 24 which may be spaced from one another in a circumferential direction such that a portion of the strip 12 may not be covered by a groove 24 and may be exposed or visible between neighboring, i.e. adjacent grooves 24. As such, the strip 12 can be visibly inspected to confirm that the strip 12 remains properly attached to the skin 108 of the fuselage sections 104 Alternatively, the plurality of slots 24 may have them supported against each other in a circumferential direction such that the entire strip 12 can be covered by the slots 24.
    [0042] In the illustrated embodiment, each socket 24 may have a "U"-shaped channel or cross-section that includes the base 32, formed from the tapered section 34 and the flat section 36, and opposing upright edges 42 (individually identified as a first upright edge 42a and a second upright edge 42b). A first upright edge 42a is positioned toward a first longitudinal side 30a and a second upright edge 42b is positioned toward a second longitudinal side 30b. In other embodiments, the sockets 24 may have other cross-sectional shapes, including "C" shapes, "L" shapes, inverted "Pi" shapes, and flat shapes.
    [0043] The upright edges 42 can add rigidity to the grooves 24 and may be positioned close to the raised portions 116 of the gussets 114 depending on the dimensions of the flat section 36 of the groove 24 and the spaced distance of the gussets 114. upright edges 42 can increase the stability of the splice joint 106, especially under compressive loads.
    [0044] If a segmented strip 12 is used, then the sockets 24 can also be used as splicing plates for the adjacent strip segments. An advantage of the disclosed splice joint 106 is that the ends of the gussets 114 are left open, which allows moisture caused by condensation and other sources to escape the gussets 114 for sufficient drainage.
    [0045] Yet another advantage of the disclosed splice 10 is that the raised portions 116 of the opposing stiffening stringers 114 are not joined through the splice joint 106, which makes the recesses 24 relatively easy to install due to the fact that the raised portions 116 don't have to be in perfect alignment. Additionally, the ability to change or adjust the position of the socket 24 along both sides 16 of the strip 12 allows the longitudinal elements 38, 40 to receive the raised portion 116 of the stiffener 114 if not perfectly aligned with a stiffener stringer. opposite 114.
    [0046] The splice 10 may optionally include a frame 44 that extends circumferentially around at least a portion of the splice joint 106 between adjacent fuselage sections 104. The frame 44 may generally be positioned to cover the intermediate region 20 of strip 12 between longitudinally opposed recesses 24. In the illustrated embodiment, frame 44 may be of a generally "Z" shape configured to have an upwardly facing lower flange 46 to be secured to strip 12. Lower flange 46 may be The upper flange 48 may be segmented into feet in a manner similar to that described for lower flange 46. Although a Z-shaped frame 44 is shown, those skilled in the art will note that any suitable frame can be used, such as a C-shape frame, an I-shape frame, a J-shape frame or the like.
    [0047] The frame 44 may also be formed of a composite material, such as CFRP or similar material, such as the same composite that forms the strip 12 or the sockets 24. Alternatively, the frame 44 may be formed of a metallic or other material. , such as aluminum or titanium.
    [0048] While the disclosed splice 10 is illustrated as being constructed of numerous separate parts (e.g., strip 12, fittings 24, frame 44), in other embodiments, two or more of these parts may be integrated into a single part that performs the function or has the characteristics of two or more parts.
    [0049] Referring next to Figure 9, one benefit of splice 10 disclosed is the contemplated elimination of the need for metallic components, such as titanium metal components, at splice joint 106 and the significant reduction (if not elimination complete) of the need for shims or non-conformity labels during assembly of the fuselage 102. A further benefit of the splice 10 disclosed is the significant reduction in manufacturing flow times of the fuselage assembly 102 contemplated when installing, i.e., fastening, the splicing strip 12 and a part of the plurality of splicing slots 24a to the first fuselage section 104a prior to the integration of the adjacent second fuselage section 104b and the corresponding part of the plurality of longitudinally opposed splicing slots 24b. Optionally, the frame 44 can also be installed prior to the integration of the second fuselage section 104b.
    [0050] As best illustrated in Figures 4, 6 and 8, the interface between the tapered region 22 of the strip 12 and the tapered sections 34 of the recesses 24 allows the recesses 24 to be secured to the strip 12 and the panel 110 without requiring placement of shims and allows for location tolerances between adjacent fuselage sections 104 and sockets 24. For example, during installation of sockets 24, each socket 24 may be peeled longitudinally forwards or backwards, i.e. towards the bow and aft, with respect to strip 12 to account for varying thicknesses and termination locations of flange portions 118 of stiffener spars 114 to provide a shimless splice joint 106. Opposite tapered surfaces of strip 12 and recesses 24 can also minimize center of gravity shift across the splice joint 106, which can reduce airplane loads along the splice joint 106. Additionally, the tapered interface can also decrease r transferring load on fasteners and more evenly transferring load forces between splicing elements, i.e. strip 12, fittings 24, and designing joined fuselage sections 104.
    [0051] Referring to Figure 10, the splicing strip 12 can be formed through the dripping compound layer to internally reach the tapered regions 22 without the requirement of post forming machining. The splice fittings 24 can also be formed through the layer of dripping compound to internally reach the tapered section 34 without the requirement for post forming machining. Strip 12 and inserts 24 may include an internal tapered structure 50 surrounded by a plurality of fiber layers and a cured resin, i.e., composite layer 52, to form a composite structure. For example, the strip 12 may include an outer continuous fiber region and an inner tapered fiber region. A plurality of layers may accumulate to form a lower continuous fiber zone. A further plurality of layers may accumulate at a perimeter boundary of the lower continuous fiber zone to form the tapered fiber zone or the tapered structure 50. Alternatively, the tapered structure 50 may be formed from alternative materials. A further plurality of layers may be built up over the tapered fiber zone to form an upper continuous fiber zone. Layers 52 may be prepreg composite layers. Splice fittings 24 may be formed by a substantially similar process to form tapered section 34 of base 32.
    [0052] An advantage of using composite materials instead of metals is that the strip 12, fittings 24 and underlying panels 110 (e.g. skins 108 and reinforcing stringers 114) can have coefficients of thermal expansion at least generally similar. As a result, temperature fluctuations experienced during the operation of the aircraft 100 may not cause uneven thermal expansion between the splice 10 and the underlying panels 110, which may not induce significant stresses on the splice joint 106. Another advantage of using composite materials is the elimination of the need for insulation of different material.
    [0053] One skilled in the art can appreciate that in addition to metallic composites and materials, in still other embodiments, the panels 110, strip 12 and fittings 24, and combinations thereof, may include other materials, including hybrid materials such as laminated fiber/metal, including fiberglass/aluminium laminates and titanium-reinforced graphite (Ti/Gr) laminates.
    [0054] The splice 10 may include a plurality of fasteners that join the splice components 10 to one another and to fuselage sections 104. Various fastener patterns to which fasteners can be installed in order to connect the splice 10 to the sections fuselage 102 may be used and have not been illustrated. Fasteners may extend through strip 12; of the frame, such as around the lower flange 46; and the fittings 24, such as both the tapered section 34 and the longitudinally extending elements 38, 40 of the flat section 36 so as to connect the components to each other and to the fuselage sections 104.
    [0055] In order to install the plurality of fasteners, a plurality of holes must be formed, as by means of drilling, in the various components of the splice 10 to receive the respective fasteners. Since a number of holes, and in certain embodiments a majority of holes, can be formed through components, for example, strip 12, fittings 24 and optionally frame 44, which are formed from a composite material such as CFRP, Holes can be formed more quickly and efficiently than comparable hole formation through titanium components or other metal components. Additionally, holes that are formed through the composite components of a splice 10 can be formed even more efficiently and more cost-effectively than comparable holes formed through components of comparable metal or titanium, since the components of amendment 10 do not need to be separated and deburred.
    [0056] Another advantage of the disclosed 10 splice is that components can be fastened with titanium fasteners, as opposed to Inconel® fasteners or other steel fasteners, as composite components may not be as sensitive to fatigue of the joint. same way as for comparable metal parts. As such, composite components generally do not require the high level of gripping forces required by metal parts using Inconel® fasteners or other steel fasteners.
    [0057] Although various embodiments of the revealed splicing joint have been shown and described, modifications may occur to those skilled in the art upon reading the specification. The present application includes such modifications and is limited only by the scope of the claims.
    [0058] In accordance with one aspect of the present description, a seam is provided for joining adjacent panels of a carcass structure comprising: a strip comprising a first side, a second side, a generally flat intermediate region, a first region tapered region extending downwardly near said intermediate region to said next first side, and a second tapered region extending downwardly near said intermediate region to said second second side, a first socket comprising a tapered section and a flat section, said tapered section is secured to said first tapered region of said strip; and, a second socket comprising a tapered section and a flat section, said tapered section being secured to said second tapered region of said strip. Advantageously, the splice additionally comprises a frame secured to said intermediate region of said strip. Advantageously, the splice comprises a composite material. Advantageously, the strip further comprises a tapered internal structure surrounded by a carbon fiber reinforced plastic.
    [0059] Furthermore, the present invention may comprise modalities according to the following clauses:
    [0060] Clause 1. A splice of the housing structure comprising: a first panel comprising a first edge; a strip secured to said first panel and extending beyond said first edge, said strip comprising a first tapered region; and, a first socket comprising a tapered section and a flat section, said tapered section being secured to said first tapered region of said strip.
    [0061] Clause 2. The seam of Clause 1, wherein at least a portion of said first tapered region of said strip is positioned between said first mortise and said first panel.
    [0062] Clause 3. The seam of Clause 1, further comprising: a second panel comprising a second edge, the second edge being positioned in edge direction alignment with said first edge to form a seam joint ; said strip bridges said splice joint and is secured to said second panel, said strip further comprising a second tapered region; and, a second socket comprising a tapered section and a flat section, said tapered section being secured to said second tapered region of said strip.
    [0063] Clause 4. The seam of Clause 3, wherein at least a portion of said second tapered region of said strip is positioned between said second groove and said second panel.
    [0064] Clause 5. The amendment of Clause 3, wherein said first panel and said second panel are fuselage sections of an aircraft.
    [0065] Clause 6. The amendment of Clause 3, wherein said strip comprises a top surface, a bottom surface, a first side, and a second side, said top surface comprising a generally flat intermediate region ; wherein said first tapered region extends downwardly near said intermediate region to said next first side; and, wherein said second tapered region extends downwardly near said intermediate region to said next second side.
    [0066] Clause 7. The seam of Clause 6, wherein said flat section of said first groove extends beyond said first side of said strip and is secured to said first panel and said flat section of said second groove extends beyond said second side of said strip and is secured to said second panel.
    [0067] Clause 8. The amendment of Clause 7, wherein said flat section of said first socket and said second socket each comprises a first longitudinal element and a second spaced longitudinal element.
    [0068] Clause 9. The amendment of Clause 1, wherein said strip comprises a composite material.
    [0069] Clause 10. The amendment of Clause 9, wherein said strip further comprises a tapered internal structure surrounded by a carbon fiber reinforced plastic.
    [0070] Clause 11. The amendment of Clause 6, which additionally comprises a frame attached to said intermediate region of said strip.
    [0071] Clause 12. A seam for joining adjacent panels of a carcass structure comprising: a strip comprising a first side, a second side, a generally flat intermediate region, a first tapered region extending downwardly near said intermediate region to the next said first side, and a second tapered region extending downwardly near the said intermediate region to the next said second side, a first socket comprising a tapered section and a flat section, said tapered section is secured to said first tapered region of said strip; and, a second socket comprising a tapered section and a flat section, said tapered section being secured to said second tapered region of said strip.
    [0072] Clause 13. The amendment of Clause 12, which further comprises a frame attached to said intermediate region of said strip.
    [0073] Clause 14. The amendment of Clause 12, wherein said strip comprises a composite material.
    [0074] Clause 15. The amendment of Clause 14, wherein said strip further comprises a tapered internal structure surrounded by a carbon fiber reinforced plastic.
    [0075] Clause 16. A method of joining a carcass structure, said method comprising the steps of: providing a first panel of said carcass structure comprising a first edge; providing a strip comprising a first tapered region and a second tapered region; positioning said strip adjacent to said first edge such that said first tapered region is adjacent to said first panel and said second tapered region extends away from said first edge; securing said strip to said first panel; providing a first socket comprising a tapered section and a flat section; positioning said first socket such that said tapered section is adjacent to said first tapered region of said strip and said flat section is adjacent to said first panel; securing said first socket to said strip; and securing said first socket to said first panel.
    [0076] Clause 17. The method of Clause 16, which further comprises the steps of: providing a second panel of said carcass structure comprising a second edge; positioning said second panel adjacent to said first panel such that said first and second edges are in alignment towards the edge forming a seam joint and said second tapered section is adjacent to said second panel and said strip bridges it. said splice joint; securing said strip to said second panel; providing a second socket comprising a tapered section and a flat section; positioning said second groove such that said tapered section is adjacent to said second tapered region of said strip and said flat section is adjacent to said second panel; securing said second socket to said strip; and securing said second socket to said second panel.
    [0077] Clause 18. The method of Clause 17, which additionally comprises the steps of: providing a frame; positioning said frame adjacent to said strip between said first tapered region and said second tapered region; and attaching said frame to said strap.
    [0078] Clause 19. The method of Clause 18, wherein said first panel and said second panel are fuselage sections of an aircraft.
    [0079] Clause 20. The method of Clause 18, wherein at least one of said strip, said first socket, said second socket and said frame comprises a composite material.
    权利要求:
    Claims (11)
    [0001]
    1. Seam of the housing structure (10) characterized in that it comprises: a first panel (110a) comprising a first edge (120a) and a first reinforcing stringer (114) spaced from the first edge (120a); a second panel (110b) comprising a second edge (120b) and a second reinforcing spar (114) spaced from the second edge (120b), said second edge (120b) being positioned in alignment towards the edge with the said first edge (120a) of the first panel (110a) to form a seam joint (106); a strip (12) secured to said first panel (110a) and to said second panel (110b), such that said strip (12) bridges the splice joint (106), said strip (12) comprises a top surface (18), a flat bottom surface (14), a first side (16a) and a second side (16b), wherein said flat bottom surface makes contact with said first panel (110a) and said second panel (110b) and wherein said upper surface comprises a flat intermediate region (20), a first tapered region (22a) extending downwardly from near said intermediate region to near said first side, and a second tapered region (22b) extending downwardly from proximate said intermediate region to proximate said second side; a first socket (24a) comprising a first tapered section (34) and a first flat section (36), said first tapered section (34) is secured to said first tapered region of said strip (12), and said first flat section (36) extends beyond said first side of said strip (12) and attached to said first panel (110a); and a second socket (24b) comprising a second tapered section (34) and a second flat section (36), said second tapered section (34) is secured to said second tapered region of said strip (12), and said second flat section (36) extending beyond said second side of said strip (12) and secured to said second panel (110b).
    [0002]
    2. A splice according to claim 1, characterized in that at least a part of said first tapered region (22a) of said strip (12) is positioned between said first tapered section (34) of said first socket and said first panel (110a).
    [0003]
    3. Amendment, according to claim 1, characterized in that said first panel (110a) and said second panel are fuselage sections of an aircraft.
    [0004]
    4. A splice according to claim 1, characterized in that said first flat section (36) of said first socket (24a) and said second flat section (36) of said second socket (24b) each , comprises a first longitudinal element (38) and a spaced apart second longitudinal element (40), wherein a portion of said first reinforcing stringer (114) fits between said first longitudinal element and said second longitudinal element of said first flat section , and wherein a portion of said second longitudinal spar (114) fits between said first longitudinal element and said second longitudinal element of said second flat section.
    [0005]
    5. Amendment, according to any one of the preceding claims, characterized in that said strip (12) comprises a composite material.
    [0006]
    6. Splice, according to any one of the preceding claims, characterized in that said strip (12) additionally comprises a tapered internal structure surrounded by a carbon fiber reinforced plastic.
    [0007]
    A splice, according to claim 1, characterized in that it additionally comprises a frame (44) attached to said flat intermediate region of said strip (12) and arranged between said first socket (24a) and said second socket (24b).
    [0008]
    8. Method of joining a carcass structure, said method being characterized in that it comprises the steps of: providing a first panel (110a) comprising a first reinforcing stringer (114) and a second panel (110b) comprising a second reinforcing stringer (114) of said carcass structure in alignment in the edge direction to form a seam joint (106); providing a strip (12) for bridging said splice joint (106), said strip (12) comprising a top surface (18), a flat bottom surface (14), a first side (16a) and a second side (16b), wherein said flat bottom surface makes contact with said first panel (110a) and said second panel (110b), and wherein said top surface comprises a flat intermediate region (20). ), a first tapered region (22a) extending from near the intermediate region to near the first side, and a second tapered region (22b) extending from near the intermediate region to near said second side; securing said strip (12) to said first panel (110a) and said second panel (110b); positioning a first socket (24a) comprising a first tapered section (36) and a first flat section (36) such that said first tapered section (36) is adjacent to said first tapered region (22a) of said strip (12) and said first flat section (36) extends beyond the first side of said strip and is adjacent to said first panel (110a); securing said first tapered section of said first socket (24a) to said first tapered region of said strip (12); securing said first flat section (36) of said first socket (24a) to said first panel (110a); positioning an adjacent second panel (24b) comprising a second tapered section (34) and a second flat section (36) such that the second tapered section is adjacent to said second tapered region of said strip and said second flat section extends beyond said second side of said strip and adjacent to said second panel; securing said second tapered section of said second socket (24b) to said second tapered region of said strip; and securing said second flat section of said second socket (24b) to said second panel (110b).
    [0009]
    Method according to claim 8, characterized in that it additionally comprises the steps of: positioning a frame (44) in said flat intermediate region of said strip (12) between said first socket (24a) and said second socket (24b); and securing said frame (44) to said flat intermediate region of said strip (12).
    [0010]
    10. Method according to claim 9, characterized in that said first panel (110a) and said second panel are fuselage sections of an aircraft.
    [0011]
    Method according to claim 9 or 10, characterized in that at least one of said strip (12), said first socket (24a), said second socket (24b) and said frame (44) ) comprises a composite material.
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    同族专利:
    公开号 | 公开日
    KR101858274B1|2018-05-15|
    ES2622301T3|2017-07-06|
    EP2727821B1|2017-03-29|
    BR102013027906A2|2014-10-21|
    AU2013228054B2|2016-12-22|
    CN103786868B|2017-09-12|
    RU2658211C2|2018-06-19|
    JP6342641B2|2018-06-13|
    US20140117157A1|2014-05-01|
    AU2013228054A1|2014-05-15|
    CA2828723C|2016-11-22|
    CN103786868A|2014-05-14|
    EP2727821A1|2014-05-07|
    KR20140055989A|2014-05-09|
    JP2014111433A|2014-06-19|
    RU2013148371A|2015-05-10|
    US8960606B2|2015-02-24|
    CA2828723A1|2014-04-30|
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    法律状态:
    2014-10-21| B03A| Publication of a patent application or of a certificate of addition of invention [chapter 3.1 patent gazette]|
    2018-11-21| B06F| Objections, documents and/or translations needed after an examination request according [chapter 6.6 patent gazette]|
    2020-02-27| B06U| Preliminary requirement: requests with searches performed by other patent offices: procedure suspended [chapter 6.21 patent gazette]|
    2021-11-09| B09A| Decision: intention to grant [chapter 9.1 patent gazette]|
    2022-01-11| B16A| Patent or certificate of addition of invention granted [chapter 16.1 patent gazette]|Free format text: PRAZO DE VALIDADE: 20 (VINTE) ANOS CONTADOS A PARTIR DE 30/10/2013, OBSERVADAS AS CONDICOES LEGAIS. |
    优先权:
    申请号 | 申请日 | 专利标题
    US13/665,664|US8960606B2|2012-10-31|2012-10-31|Circumference splice for joining shell structures|
    US13/665,664|2012-10-31|
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