• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
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    • 专利摘要:
    vertically integrated stiffening spars an apparatus and processes provide for the reinforcement of various components of an aircraft using vertically oriented circumferential spars. in accordance with the embodiments described in the present specification, the tubular metallic skin of an aircraft fuselage can be reinforced by using reinforcing spars, which are oriented vertically and circumferentially. according to other embodiments, the wing ribs of a wing may be aligned with the circumferential ribs. according to still other embodiments, the wing may also have one or more wing spars having an elliptical aperture. the elliptical shape can be configured for attachment to a fuselage so that an outer surface of the opening is disposed close to an inner surface of the fuselage.
    公开号:BR102013030182B1
    申请号:R102013030182-5
    申请日:2013-11-25
    公开日:2022-01-18
    发明作者:John H. Moselage, Iii
    申请人:The Boeing Company；
    IPC主号:
    专利说明:

    BACKGROUND
    [001] An aircraft commonly uses thin strips of material, or "reinforcement spars", oriented longitudinally parallel to a central axis of the aircraft, to provide mechanical strength and rigidity characteristics to the metallic coating covering the fuselage. Existing aircraft fuselages are typically manufactured in cylindrical sections, which are joined end to end to create a fuselage having the desired length. In the fabrication of a fuselage section, several reinforcing spars are bonded to the metal skin, so that each reinforcing spar terminates at the front and rear ends of the fuselage section. When fuselage sections are joined to adjacent fuselage sections, each end of a stiffening spar must be spliced into a corresponding end of a stiffening spar of the adjacent fuselage section. This splicing operation is a labor intensive, tedious process.
    [002] In addition, with each seam in a conventional fuselage configuration, there is a potential for delamination or other structural failure or fatigue, due to the common loads and moments induced by the aircraft fuselage during flight. To prevent these structural failures, splicing plates and fasteners are typically used, which significantly increases the weight and cost of the aircraft.
    [003] It is in relation to these considerations and others that the invention described in this descriptive report is presented. SUMMARY
    [004] It should be noted that this summary is provided to introduce a selection of concepts, in a simplified form, which are further described below in the detailed description. This summary is not intended to be used to limit the scope of the claimed topic.
    [005] The apparatus and processes provide a reinforced aircraft fuselage, which uses vertically oriented, circumferential reinforcement spars. In accordance with an aspect of the invention provided in the present specification, an aircraft fuselage may include a metallic skin, having an internal surface. The metallic coating may extend along a longitudinal axis. A number of reinforcing stringers may be arranged substantially normal to the longitudinal axis. The various reinforcing stringers can be oriented vertically and circumferentially along the longitudinal axis.
    [006] According to another aspect, an aircraft may have a fuselage, which may include a cylindrical metallic skin, having an inner surface. The cylindrical metallic shell may extend longitudinally along a first axis. A number of cylindrical stiffeners may be arranged substantially normal to the first axis. Cylindrical stiffening stringers may be oriented vertically circumferentially along the first axis. The aircraft may also have a wing with a metallic wing coating. The metallic wing skin may have an inner surface, with a first end adjacent to the fuselage and a second end distant from the first end. The wing may have multiple wing reinforcement spars attached to the inner surface of the metallic wing skin. The wing stiffener spars may be oriented so that they substantially align with the circumferential spars of the fuselage.
    [007] According to another aspect, a process for reinforcing an aircraft may include forming a cylindrical metallic skin, having an inner surface that extends longitudinally along a first axis. Several reinforcing stringers are formed which are circumferentially oriented and substantially parallel to each other. The reinforcing stringers are coupled to the cylindrical metal skin so that at least a part of the stringers is aligned normal to the first axis.
    [008] The features, functions and advantages, which have been discussed, can be obtained independently in various embodiments of the present invention, or can be combined in more other embodiments, the additional details of which can be seen with reference to the description presented below and the drawings. BRIEF DESCRIPTION OF THE DRAWINGS
    [009] Figure 1 is a top perspective view of an exemplary fuselage and wing sections of an aircraft using circumferential stiffening spars, according to the various embodiments presented in the present specification.
    [0010] Figure 2 is a top perspective view illustrating affixation of adjacent fuselage sections, metallic skins and circumferential reinforcement spars, in accordance with the various embodiments presented in the present specification.
    [0011] Figure 2A is a top perspective view, in the foreground illustrating the attachment of an aircraft metallic skin to a fuselage section, according to the various embodiments presented in this specification.
    [0012] Figure 2B is a top perspective, foreground view illustrating the coupling of adjacent fuselage sections, in accordance with the various embodiments presented in the present specification.
    [0013] Figures 3, 3A and 3B are various views of exemplary circumferential blade reinforcement spars, according to the various embodiments presented in the present specification.
    [0014] Figures 4, 4A and 4B are various views of exemplary circumferential corrugated reinforcement stringers, in accordance with the various embodiments presented in the present specification.
    [0015] Figures 5 and 5A are several views illustrating the attachment of wing reinforcement spars to the fuselage of an aircraft, according to the various embodiments presented in the present specification.
    [0016] Figures 6A and 6B are perspective views illustrating wing spar and circumferential stiffening spar configurations, in accordance with the various embodiments presented in the present specification.
    [0017] Figure 7 is a front end view in perspective of an aircraft fuselage and wing, showing the exemplary force vectors, according to the various embodiments presented in the present specification.
    [0018] Figure 8 is a process flow diagram illustrating a process for using circumferential stiffening spars, and a process for attaching a wing to a fuselage using wing spars, in accordance with the various embodiments presented in this specification. . DETAILED DESCRIPTION
    [0019] The detailed description presented below provides aircraft reinforcement structures. As briefly discussed above, an aircraft commonly uses conventional stiffening spars, which are oriented longitudinally parallel to a central axis of the aircraft, to provide mechanical strength and reinforcement characteristics to the metallic coating covering the fuselage. The fuselage of a conventional aircraft is typically formed of multiple sections, which are joined together. By using conventional stiffening stringers, each end of each stiffening stringer has to be spliced into a corresponding end of an adjacent stiffening stringer of the adjacent fuselage section. This is a labor intensive, time intensive process. Other limitations in using conventional stiffening stringers are also briefly discussed above.
    [0020] By using the concepts described in this descriptive report, the metallic cladding of an aircraft fuselage can be structurally reinforced by using reinforcing spars, which are oriented vertically and circumferentially along the fuselage. Instead of using stiffener spars, which extend along the fuselage front to back, stiffener spars, which are oriented vertically and circumferentially along the fuselage, can reduce the labor and time costs associated with spars. reinforcement that extend along the fuselage. Furthermore, in some embodiments, the reinforcing spars and/or the fuselage or metallic wing skin can be formed using various manufacturing techniques, such as a mandrel forming, which can reduce manufacturing time and cost. Exemplary techniques are described in the copending patent application entitled "Multi-Box Wing Spar and Skin", filed November 26, 2012, which is thus fully incorporated herein.
    [0021] Also, as will be described below, wing root loads can be reacted circumferentially around the fuselage. Aircraft wings can be attached to the fuselage by means of wing spars, which encircle and/or encircle a portion of the fuselage. Wing spars can be attached directly to the fuselage. The wing stiffening spars may be arranged with and secured to vertically oriented circumferential spars of the fuselage. Fuselage stiffener spars, which are configured to mate (join) with wing stiffener spars, can be sized to accommodate loads from the wing stiffener spars. Loads from the wing stiffener spars can then travel to the stiffener spars / fuselage frames to be neutralized / reacted with loads from the opposite wing at the crown and/or keel locations. It should be appreciated that the concepts described in this specification relating to an aircraft wing may also be used for other aircraft components, such as a vertical or horizontal stabilizer, without departing from the scope of this invention and the associated claims. .
    [0022] In the detailed description presented below, references are made to the attached drawings, which form a part thereof, and which are shown by way of illustration, specific embodiments or examples. With reference then to the drawings, in which like numerals represent similar elements throughout the various figures, the use of circumferential stiffening stringers will be described.
    [0023] Returning to figure 1, a top perspective view of a part of an aircraft 100 is shown. The aircraft 100 may have a fuselage 102 extending along the longitudinal axis "XY" from the front 102 of the fuselage to the rear 106 of the fuselage. Aircraft 100 may also have wings 108a and 108b. It should be understood that wings 108a and 108b are shown in a straight wing formation for illustrative purposes only. While the various embodiments may be discussed and/or illustrated with respect to wings 108a and 108b in a straight wing configuration, the concepts and embodiments described in this specification may be equally applicable to other wing configurations, including but not limited to a, an arrow wing configuration (as illustrated by way of example in Figure 6), without departing from the scope of this invention and the appended claims.
    [0024] According to an illustrative embodiment, the fuselage 102 may be reinforced along the XY axis by use of various beams, including but not limited to the keel beam 110 and/or crown beam 210 (shown in Figure 2). The fuselage 102 may also have additional structural elements, such as a support element 113, to provide further structural support and/or provide attachment of various aircraft components, such as a floor. In a conventional aircraft, several longitudinal stiffening spars, extending along the XY axis, can be used to structurally reinforce a fuselage and a metal skin (not shown). However, the embodiments described in the present specification utilize stiffening spars that are oriented vertically circumferentially along the fuselage, which reduces or eliminates the various conventional longitudinal spars used.
    [0025] The fuselage 102 has circumferential rib spars 112 which are oriented vertically along the fuselage 102. Exemplary circumferential rib spars 112a - c of circumferential rib spars 112 are shown by way of example to illustrate the orientation of the spars of reinforcement along the fuselage 102. It should be noted that the circumferential reinforcement spars 112a - c are specifically indicated and identified in Figure 1 for descriptive purposes only. It should be noted that a fuselage such as the fuselage 102 illustrated in Fig. 1 may have more or less reinforcing spars than those illustrated in Fig. 1. Furthermore, it should be borne in mind that the distance between the spars circumferential ribs 112 is illustrative only and does not reflect an intention to limit the scope of the present invention or the appended claims to equally spaced ribs, as the spatial configurations of the ribs may vary from aircraft to aircraft, or across of a fuselage.
    [0026] As previously indicated, the circumferential ribs 112 are oriented vertically circumferentially along the fuselage 102. As illustrated, the ribs 112a - c are aligned circumferentially (i.e., at least partially encircle an internal, longitudinal axis of the fuselage). 102) and vertically oriented (i.e., when viewed from the side of the fuselage 102, whereby the XY axis is horizontal from an observer's left side to an observer's right side, the circumferential stiffening stringers 112 run vertically through, or normal to the XY axis). It should be noted that the circumferential gussets 112 are illustrated as being completely vertically oriented for illustrative purposes only, and do not reflect an intention to limit the scope of the present invention or the accompanying drawings to gussets that are perfectly vertical, as some circumferential stiffening spars 112 may also be partially or substantially vertical, depending on the particular aircraft design criteria. In other words, some of the circumferential stiffening stringers 112 may have an angular displacement greater than or less than normal to the XY axis. For example, and not by way of limitation, a portion of the circumferential stiffeners 112 may be angled towards one end of the fuselage 102. In this configuration, the portion of the circumferential gussets 112, which are angled towards one end of the fuselage 102, can be angled so that other components of an aircraft can be integrated with the part of circumferential stiffening spars 112, which are inclined towards one end of the fuselage 102. For example, if wings 108a and 108b are in an arrow wing configuration, the portion of the circumferential stiffener spars 112, which are angled towards one end of the fuselage 102, can be angled so that the wing stiffener spars 114 can be more easily integrated into the fuselage 102. In some embodiments, the circumferential stiffening spars 112 are constructed using metal forming processes, or, if produced made of a polymeric, composite or plastic material, may be molded. Examples of various composite, plastic or other materials may include carbon fiber reinforced composites or carbon fiber reinforced thermoplastic composites.
    [0027] The wings 108a and 108b can be integrated with the fuselage 102 through the use of wing spars (examples of which are described in the present specification below) and wing reinforcement spars 114 and 116, respectively. Specific examples of wing spar 114 are identified as wing spars 114a and 114b, and wing spar 116 is identified as wing spars 116a and 116b. Wing reinforcement spars 114 and 116 can be used not only to reinforce the various aspects of wings 108a and 108b, but can also be used to secure wings 108a and 108b to fuselage 102 and/or transfer flight loads from wings 108a and 108b for fuselage 102. It should be noted that a wing, such as wing 108a or 108b, may have more or less stiffening spars than those illustrated in figure 1. Furthermore, it should be considered that the distance between the wing stiffeners 114 or 116, or their relative angular positions to one another, are illustrative only and do not reflect an intention to limit the scope of the present invention or claims associated with parallel, equally spaced stiffeners, as the Spatial configurations of reinforcement spars may vary from aircraft to aircraft or across a wing.
    [0028] Figure 2 illustrates how the metallic skin 204 of an aircraft may be coupled to the fuselage 102. As shown in Figure 2, the fuselage 102 has exemplary circumferential stiffening spars 112. It should be noted that the fuselage 102 may be constructed by use more or less of the circumferential stiffening spars 112, as various fuselages may use fewer or more stiffening spars, depending on the particular fuselage section. The circumferential stiffening stringers 112 are illustrated as being circumferentially oriented vertically around a central axis, thereby creating a column or cylinder. It should be noted that the present invention is not limited to any particular way of coupling. For example, the metal skin 204 can be arranged with the fuselage 102 extended, or the metal skin 204 can be attached to the fuselage 102 after the fuselage 102 has been disposed.
    [0029] The fuselage section 202 has metallic skin 204 disposed thereon, having a first end 206 and a second end 208. The metallic skin 204 can be coupled to the fuselage section 202 by use of various means. For example, and not by way of limitation, the first end 206 can be heat welded to the second end 208, if the metallic coating 204 is formed of a heat weld material. In another example, the first end 206 may be welded, riveted, or otherwise secured to the keel beam 110. In some embodiments, the metal cladding 204 may be coupled using various coupling processes to couple the metallic cladding 204 to the fuselage section 202. The manner in which components can be secured may depend on various factors such as, but not limited to, the type of material used for the circumferential stiffening spars 112, the structural load requirements of the aircraft, etc. . It should also be considered that the metallic skin 204 may be coupled to other components of the aircraft, such as the crown beam 210 or the side beam 212. Furthermore, it should be considered that the metallic skin 204 can be secured in more than one way. structural component of fuselage section 202.
    [0030] As in more conventional aircraft designs, fuselage 102 can have multiple sections coupled together. By way of example, the fuselage 102 of Figure 2 is shown having three exemplary sections: the fuselage section 202; the wing section 214 and the aft section 216. Further description of how the wing reinforcement spars 116 can be attached to the fuselage 102, in the wing section 214, can be found in more detail below, and by way of example , in Figures 6A and 6B. It should be noted that the present invention is not limited to any particular fuselage configuration, as the number and type of fuselage sections may vary depending on aircraft design or other design factors.
    [0031] By using various embodiments of the present invention, labor costs and/or time can be reduced when attaching one fuselage section 102 to another section. In aircraft construction using conventional longitudinal stiffeners, connecting one fuselage section to another fuselage section may require aligning up to one hundred parts, including the longitudinal stiffening spars in one section, with their counterparts in a section to be united, adjacent. A significant amount of time is spent lining up two sections and drilling precise holes to join the reinforcing stringers to their counterparts. After the sections are aligned and holes drilled, the sections are simply disassembled to deburr the various parts. Then, the reinforcing stringers are secured together using conventional fastening means such as, but not limited to, rivets, screws or fasteners. In some construction processes, it may take 10 working days to fix one section to the other. Furthermore, because rivets, spacers, and other parts are typically used to join together reinforcing spars by means of two or more cylindrical sections, the weight of the aircraft can be adversely affected.
    [0032] According to one embodiment, when using circumferential stiffening stringers 112, the fuselage sections may be joined using, by way of example and not by way of limitation, an Inner Mold Line (IML) splicing strip. , as illustrated in Figure 2B. The fuselage section 216 can be supported on the adjacent fuselage section 218, shown in more detail in Figure 2B. Because the circumferential stiffeners 112 are arranged around a central axis of the fuselage 102, the circumferential stiffeners 112 do not need to be aligned longitudinally along the central axis, only a few distinct beams need to be aligned. Once the aircraft manufacturer rotatably aligns the fuselage section 216 and the fuselage section 218, the manufacturer can couple the two sections together using the IML splicing strip 220. The splicing strip 220 can be used to couple the section of fuselage 216 in fuselage section 218 using various processes including, but not limited to, fasteners, binders and heat welds. It should be understood that the use of IML 220 splicing strip is for descriptive purposes only and is not intended to limit the scope of the present invention or associated claims, as other types of strip may be used. Further, it may be noted that a splicing strip, such as an IML 220 splicing strip, may be constructed of various materials, and may be joined to one or more parts of the aircraft, using various fastening processes, such as fasteners. , chemical bonding or thermal bonding.
    [0033] Various types of stiffening stringers can be used in conjunction with the various embodiments of the present invention. Two primary embodiments of stiffener stringer configurations will be described in the present specification, although other circumferential stiffener stringer configurations will be considered. The first exemplary circumferential reinforcement spar embodiment is illustrated in Figures 3, 3A and 3B. In Figure 3, the fuselage 302 is shown constructed using blade reinforcement spars 312, illustrated in more detail in Figure 3A.
    [0034] Figure 3A shows a small part of the fuselage 302 with the metal fuselage skin 304 cut away for clarity. The fuselage 302 is shown constructed using multiple circumferential blade reinforcement spars, illustrated by way of example as circumferential blade reinforcement spars 312a - c. The circumferential blade reinforcement stringers 312 are circumferential stringers, and may be constructed using a variety of materials, including, but not limited to, various metals, polymers, and composites. According to various embodiments, the circumferential blade reinforcement stringers 312 can be manufactured from composite materials, such as composite materials having high axial modulus fibers for increased rigidity. Composite materials provide high mechanical strength characteristics with substantial weight savings compared to conventional metallic materials. Even more, it should be appreciated that the circumferential blade reinforcement stringers 312 can be formed in accordance with various processes, including, but not limited to, roll forming, die casting, or a forming mandrel. The circumferential blade reinforcement spars 312 may be attached to the metal skin 304 of the fuselage 302 by use of a variety of processes, including, but not limited to, thermal bonding, thermal curing, chemical bonding, or bonding.
    [0035] Figure 3B is a cross-sectional view of a circumferential blade reinforcement spar 312b, taken along line A - A in Figure 3A. The circumferential blade reinforcement spar 312b is shown having a web section 306 and a top end cap 308. The top end cap 308 may be disposed adjacent the metallic skin 304 of the aircraft. The circumferential blade reinforcement spar 312a can be made of any suitable material, depending on the particular application. In some embodiments, the circumferential blade reinforcement spar 312, web section 306 and top end cap 308 resemble, or are in, a substantially identical or similar shape to a capital letter "T".
    [0036] Figures 4, 4A and 4B illustrate a second embodiment of a circumferential reinforcement spar. In Figure 4, the fuselage 402 is constructed by using corrugated circumferential ribs 412 in place of other types of ribs, such as the circumferential blade ribs 312 of Figure 3. Figure 4A is a close-up view. of the circumferential corrugated reinforcing stringers 412 of the fuselage 402, with the metal skin 404 of the fuselage 402 cut away for clarity. In Figure 4A, exemplary circumferential corrugated ribs 412a and 412b of circumferential corrugated ribs 412 may be continuous ribs of one or more rib sections. According to various embodiments, the circumferential corrugated stiffeners 412 can be manufactured from composite materials, such as composite materials having a high axial modulus for greater rigidity. Even more, it should be appreciated that circumferential corrugated reinforcing stringers 412 can be formed according to a variety of processes, including, but not limited to, roll forming, die casting or a mandrel forming. The circumferential blade reinforcement spars 412 may be attached to the metal skin 404 of the fuselage 402 using a variety of processes, including, but not limited to, chemical bonding, thermal bonding, or bonding.
    [0037] Figure 4B is a cross-sectional view of an embodiment of the circumferential corrugated stiffeners 412a and 412b, taken along the line B - B in figure 4A. In this embodiment, rather than being configured as separate, individual gussets, such as the circumferential blade gussets 312 of Figure 3, the circumferential corrugated gussets 412 may be configured as having a series of contiguous peaks and valleys and sequential. As illustrated in Figure 4B, the circumferential corrugated stiffeners 412a and 412b are shown having top sections 414a, 416a, respectively, and bottom sections 418a and 418b, respectively. Top section 414a is illustrated as being contiguous with bottom section 418a, which is, in turn, illustrated as being contiguous with top section 414b. The top section 414b is, in turn, illustrated as being contiguous with the bottom section 418b. This pattern may be continued by the configuration of circumferential corrugated stiffeners 412, as partially illustrated at part 420.
    [0038] Various concepts and technologies described in the present specification may provide the ability to ultimately install a metallic cladding, such as the metallic cladding 404 of Figure 4. In some embodiments, this may allow the material of the metallic cladding to be a metallic cladding. of Al or Ti over a composite substructure. This can be advantageous for purposes of assistance in hail or lightning. It may also allow the metal coating to be replaced in the event of damage, where an integral composite metal coating must be patched. Additionally, having a corrugated substructure, as in the embodiment described in Figure 4, can provide a space (such as the area between the top sections 414a - b and the bottom sections 418a - b) in which through screws with protruding heads can be housed for mounting internal elements such as overhead storage receptacles, etc. Placing the metallic coating last can also provide for this type of work to be done at a more intensive pace than that of conventional techniques, by virtue of the greater access to both sides of the fastener.
    [0039] Various manufacturing techniques can be used to form the corrugated circumferential ribs 412. For example, the corrugated circumferential ribs 412 can be formed by using a thermoplastic or thermoset material in a mold or press. An example of a material that can be used is carbon fiber reinforced polymer (CFRP), but it should be considered that other materials, including metals, as well as polymeric and non-polymeric materials, can be used in accordance with the various requirements. achievements.
    [0040] As discussed with reference to Figure 1, wing reinforcement spars can be secured to a fuselage by use of attachment processes. Figures 5 and 5A illustrate an exemplary embodiment for attaching wing reinforcement spars 516 of wing 508 to fuselage 502, if wing 508 is in a straight or near straight wing configuration. Because the circumferential stiffening spars 512 can be oriented vertically along the fuselage 502, the wing stiffening spars 516 can be angularly adjusted to the fuselage 502 by virtue of their common directions.
    [0041] Figure 5A is a close-up view illustrating how the wing stiffener spars 516 may align with the circumferential stiffener spars 512. The exemplary wing stiffener spar 516a of wing stiffener spars 516 is shown coupled to the wing stiffener spar 516a. fuselage 502, between circumferential stiffening spars 512a and 512b. If circumferential corrugated gussets are used, such as the circumferential corrugated gussets 412 of Figure 4, the wing gusset 516a may be disposed within the bottom section 514 between the circumferential gussets 512a and 512b. In this embodiment, the integration of the wing stiffener spars 516 into the fuselage may have no or minimal nominal impact on the surface profile of the fuselage 502. Further, in a similar manner, the circumferential stiffener spars 512 may have spaces between the spars. of reinforcement, as shown by way of example, the circumferential blade reinforcement stringers 312 of Figure 3A. Wing stiffener spars 516 can be placed in spaces between circumferential blade stiffener spars 312 to reduce impact on the surface profile of the aircraft.
    [0042] While attaching wing reinforcement spars to an aircraft fuselage provides some structural reinforcement, conventional wings are typically attached to a wing bracket. In some embodiments in accordance with the present invention, one or more wing spars can be used to secure a wing to the fuselage, without the need for a conventional wing bracket. According to various embodiments described in the present specification, the wing reinforcement spars can be attached directly to an aircraft frame. In some embodiments, this attachment can help transfer the lifting overhead loads, through shear, to the frames, while the wing camber loads are neutralized at the wing root, providing greater separation between the upper and lower wing roots. , to react more efficiently to the bending moment, as described in more detail in Figure 7. In some embodiments, attaching one or more wings to the fuselage, in accordance with various embodiments described in the present specification, may provide additional benefits. For example, the top of the wing 508 can be secured near the top of the fuselage 502, while the bottom of the wing 508 can be secured near the bottom of the fuselage 502. This can provide more space inside the aircraft, which Can be used for cargo, fuel, gear, etc. For example, in some embodiments, the elimination of wing support can leave more cargo space, providing a rear cargo door and continuous interior space, without the need for two side cargo doors and two spaces. Other benefits can be obtained by eliminating a wing support, whereby the present invention is not limited or dependent on obtaining any specific benefit or advantage.
    [0043] Figure 6A illustrates an exemplary wing spar configuration, which can be used to attach one or more wings to a fuselage. For the purposes of only illustrating one embodiment, the wing spar attachment mechanism is described in the present specification relating to a wing, although it should be appreciated that the same attachment mechanism may be used for other aircraft components, such as , but not limited to, a horizontal or vertical stabilizer. Front spars 610a and 610b and rear spars 612a and 612b are shown. It should be appreciated that a wing, such as the wing 508 of Figure 5, may have more or fewer spars disposed thereon than the two spars, without departing from the scope of this specification and the associated claims. For example, in some embodiments, there may be a single spar, disposed at any of several locations along the wing chord (e.g., front, middle, rear). In other embodiments, there may be more than two stringers. The present invention is not limited to, or dependent on, any specific number of stringers.
    [0044] The front rails 610a and 610b have elliptical openings 614a and 614b, respectively. Rear stringers 612a and 612b have elliptical openings 616a and 616b, respectively. Depending on the angular displacement between the wing and fuselage of the aircraft, the apertures 614a - b and 616a - b may vary in circumference and shape, i.e. the foci may vary as well as the radii. For example, on a straight wing aircraft, in which the spar can be attached to the fuselage at an angle of approximately 90 degrees, openings 614a - b and/or 616a - b may be circular. In another example, such as that illustrated in Figure 6, the front spars 610a - b or the rear spars 612a - b are secured in an arrow wing configuration. In this way, the openings 614a - b and/or the openings 616a - b can be of a more oval shape, to provide an internal space in the aircraft and to be circumferentially attached to the fuselage.
    [0045] To provide more structural rigidity as well as reactions to wing root moments, by way of example and non-limiting means, the front spars 610a - b can still be attached to the rear spars 612a - b. In the exemplary embodiment shown in Figure 6A , rear spar 612b is attached to front spar 610a at joint 618, and, in a similar manner, rear spar 612a is attached to front spar 610b at joint 620.
    [0046] Figure 6B illustrates the attachment of wing spars and wing spars, according to various embodiments, to a fuselage. The fuselage section 600 has wings 602 and 604 attached thereto. As discussed previously in some embodiments, the wing stiffener spars 606 of the wing 602 and the wing stiffener spars 606 of the wing 604 may be coupled to the fuselage section 600. vis-à-vis use of crown beam 630. One or more fuselage stiffening spars, such as circumferential stiffening spar 628, may be used to provide structural support to fuselage section 600. To provide structural support and securement Wings 602 and/or 604 on fuselage section 600, as discussed by way of example in Figure 6a, wing spars may be used.
    [0047] Front spars 610a and 610b have elliptical openings 614a and 612b, and rear spars 612a and 612b have elliptical openings 616a and 616b, respectively (shown in Figure 6A), which are formed and placed in a manner that spans the section 600, i.e. provide a circumferential displacement near the outer surface of the fuselage. Thus, while providing the attachment of the wings 602 and 604 to the fuselage section 600, the front spars 610a - b and the rear spars 612a - b can also provide structural rigidity to the fuselage section 600. As discussed in Figure 6A, depending on the angular displacement, between the wing and the fuselage of the aircraft, the openings 614a - b and 616a - b can vary in circumference and shape, that is, the foci can vary as well as the radii. For example, on a straight wing aircraft, in which the spar can be attached to the fuselage at an angle of approximately 90 degrees, openings 614a - b and/or 616a - b may be circular. In another example, such as that illustrated in Figure 6, the front spars 610a - b or the rear spars 612a - b are secured in an arrow wing configuration. In this way, the openings 614a - b and/or the openings 616a - b can be of a more oval shape, to provide internal space in the aircraft and to be circumferentially attached to the fuselage.
    [0048] To aid in forming the shape of the wing 602 and, among other things, provide additional structural support, the wing 602 may have the rib 622 disposed thereon. While the present invention does not require any particular advantage or appearance, the rib 622 and the front spar 610b and/or the rear spar 612b, having one or more spars mechanically attached to one or more ribs of a wing, can help to stabilize the lead/lag camber between the front spars 610a - b and the rear spars 612a - B. Various embodiments of the present invention can also help to stabilize an upper and a lower metallic coating surface (not shown) and can aid configuration into a unitary assembly so that the assembly can act as a single unit for multiple benefits. For example, the various embodiments of the present invention can help to reduce or eliminate areas or locations where forces imposed on a wing can cause the wing to curve into a first curve mode, thereby, in some cases, reducing the probability of wing curl. It should be noted that the present invention is not limited to, or dependent on, a specific number of ribs. The number of ribs on a wing can vary depending on the specific load requirements and the construction of the specific wing. One such example is illustrated in Figure 6B. In Figure 6B, the front spar 610b is attached to the rib 622 at the first joint 620, and the rear stringer 612b is attached to the rib 622 at the second joint 624.
    [0049] Further, it may be beneficial to have the wing spar of one wing attached to the rib of an opposite wing. One such example is shown in Figure 6. In Figure 6, the rear spar 612a of the wing 604 is shown attached to the rib 622 at the junction 620. This can increase the structural stability of the aircraft, among other benefits. While the present invention is not limited to any particular benefit, nor is it based on any particular application theory, by attaching the front spar 612a to the rib 622, greater structural rigidity can be promoted without additional supporting structures, such as a support bracket. wing. Through the use of circumferentially attached spars, such as the front spar 610b and rear spar 612b, and, in some embodiments, their counterparts on an opposite wing, the use of a wing bracket can be eliminated.
    [0050] If desired or necessary, additional structural support can be provided by use of additional support structures (a third, fourth, etc. wing spar), such as a cross member 626, also secured to rib 622 at junction 620 It should be noted that one or more junctions, other than junctions 620 or 622, may be used without departing from the scope of this invention and the associated claims.
    [0051] Figure 7 is an illustration showing possible forces acting on an aircraft, using various embodiments of the present invention. The fuselage 700 has wings 702 and 704 attached to it. When in flight, the weight of the fuselage 700 and wings 702 and 704 is substantially accounted for by wings 702 and 704. Lifting the aircraft creates upward forces 706 and 708 on the wings. 702 and 704, respectively. Upward forces 706 and 708 torque the fuselage 700 in opposite directions. This torque causes tension 710 along the bottom of the fuselage 700 and compression 712 along the top of the fuselage 700. In some embodiments, because the wing spars (not shown) of the wings 702 and/or 704 are circumferentially secured to the fuselage 700, stress 710 is eliminated by compression 712, and vice versa. Thus, in some embodiments, rotational forces, such as rotational forces 714 and 718, may be partially or entirely eliminated by an equal and opposite rotational force. Therefore, in some configurations, various wing loads can be neutralized in the aircraft frame.
    [0052] Additionally, various embodiments described in the present descriptive report can provide dynamic flight load absorbing capability in a better manner than conventional aircraft construction. The aircraft experiences both loaded and unloaded conditions brought on by a number of factors, including, but not limited to, fuselage pressurization and aircraft wing flex, both in-flight and off-air. The configuration of Figure 7 may allow the aircraft to receive a dynamic flight load on several circumferential stiffening spars in a fuselage part of an aircraft. In response to receiving the dynamic flight load, tensile or compressive forces, corresponding to the dynamic flight load, may be distributed circumferentially around the fuselage part substantially normal to a longitudinal axis of the fuselage part. Circumferential gussets, such as the gussets 112 of Figure 2, can be configured to flex in a desired direction as the aircraft moves from an in-flight condition to an out-of-flight condition.
    [0053] Turning now to figure 8, an illustrative routine 800, to reinforce an aircraft, is described in detail. Unless otherwise indicated, it should be considered that more or fewer operations can be performed than those shown in the figures and described in this descriptive report. Additionally, unless otherwise indicated, these operations may be conducted in a different order than described in this specification. Further, unless otherwise noted, a particular component or feature identified in a figure is for descriptive purposes only and does not represent an intent to limit routine 800 or a particular operation to the identified component or feature.
    [0054] Routine 800 begins at operation 802, in which a metallic coating 204 is formed having an inner surface. In some embodiments, the metallic coating extends longitudinally along a first X - Y axis. From operation 802, routine 800 continues to operation 804, in which a number of circumferential gussets 112 are formed. In some embodiments, the circumferential gussets 112 are circumferentially oriented and substantially parallel to each other. In other embodiments, the circumferential gussets 112 are comprised of circumferential blade stringers 312, having a top cap 308 and a web portion 306. In other embodiments, the circumferential gussets 112 are circumferential corrugated gussets 412, having a top section 414a - b and a bottom section 418a - b.
    [0055] Routine 800 continues from operation 804 to operation 806, in which a first end of at least a portion of the circumferential gussets 112 may be secured to a second end of at least a portion of the circumferential gussets 112. Various fastening processes can be used, including, but not limited to, heat welding and riveting. It should also be understood that the present invention is not limited to circumferential gussets 112 having one end secured to the other end, or to any particular manufacturing/attachment process, as circumferential gussets 112 may be secured to various components. of an aircraft, including, but not limited to, a keel beam 110 or a crown beam 210, as described by way of example in operation 808.
    [0056] Routine 800 continues from operation 806 to operation 808, in which at least a portion of the circumferential stiffening stringers 112 are secured to a beam, such as the keel beam 110 or the crown beam 210. fasteners may be used, including, but not limited to, heat welding and rivets. In addition, it should also be understood that other beams may be used, including, but not limited to, a side beam 212.
    [0057] Routine 800 continues from operation 808 to operation 810, in which at least a part of the metallic cladding 204 is coupled to at least a part of the various circumferential reinforcement spars 112, so that at least a part of the various circumferential reinforcement stringers 112 is aligned normal to the first X-Y axis. Various fastening processes may be used, including, but not limited to, heat welding and riveting.
    [0058] In some embodiments, it may be useful to attach one or more wings (or other aircraft components) to the fuselage, prior to mating the metal cladding 204 to the circumferential stiffening spars 112. Therefore, if it is desirable or necessary to attach one or more wings , for example, wing 602, in fuselage section 600, prior to operation 810, routine 800 may continue from operation 808 to operation 812, in which several wing reinforcement spars 606 are coupled to an inner surface of a skin wing metallic. In some embodiments, the wing gussets 606 may be aligned so that they are adjacent to the circumferential gussets 112. In other embodiments, the wing gussets may be positioned between one or more of the circumferential gussets 112 For example, and not by way of limitation, one or more of the circumferential corrugated ribs 412 may be arranged in the bottom sections 418a - b, between the top sections 414a - b, if the circumferential ribs 112 are spars. circumferential corrugated ribs 412. In another example, and not by way of limitation, one or more of the wing gussets 114 or wing gussets 116 may be disposed in the spaces between the blade gussets 312 having a top end cap 308 and a web part 306, whereby the blade reinforcing stringers 312 are arranged in a space between the top end cap 308 of the circumferential blade reinforcement stringers 312.
    [0059] Routine 800 may continue from operation 812 to operation 814, in which one or more wing spars 610a, 610b, 612a and 612b, having an opening 614a, 614b, 616a and 616b, respectively, are arranged circumferentially and secured in the fuselage section 600. In some embodiments according to the present invention, one or more wing spars 610b and 612b can be used to attach a wing 602 to fuselage section 600, without the need for a conventional wing bracket. It should be noted that the same attachment operation 814 can be used for other aircraft components, such as, but not limited to, a wing and horizontal or vertical stabilizers. Depending on the angular displacement between a wing 602 and a fuselage section 600 to which the wing is to be attached, the spar opening 614a, 614b, 616a and 616b may vary in circumference and shape. It should be appreciated that the wing 602 may have more or fewer spars disposed thereon than two spars, without departing from the scope of this description and the associated claims. Routine 800 may continue from operation 814 to operation 816, in which a metallic wing skin is coupled to wing spars 610a, 610b, 612a, and 612b and to wing spars 606, if not already done. It should be noted that, in some embodiments, the metallic wing skin may be integral with, continuous with, or equal to the metallic skin covering the fuselage section 600.
    [0060] Further, the description comprises embodiments in accordance with the following clauses:
    [0061] Clause 1. An aircraft, comprising: a number of circumferential stiffening spars disposed within an aircraft fuselage and configured substantially normal to a longitudinal axis of the aircraft; and a metallic cladding coupled to the various circumferential reinforcement spars.
    [0062] Clause 2. The aircraft according to clause 1, in which at least one beam is arranged generally parallel to the longitudinal axis.
    [0063] Clause 3. The aircraft according to clause 2, wherein the beam is a keel beam or a crown beam.
    [0064] Clause 4. The aircraft according to clause 2, in which at least a part of the metallic coating is attached to the beam.
    [0065] Clause 5. The aircraft according to clause 1, wherein the metallic coating comprises a composite material.
    [0066] Clause 6. The aircraft according to clause 5, wherein the composite material comprises a carbon fiber reinforced composite or a carbon fiber reinforced thermoplastic composite.
    [0067] Clause 7. The aircraft according to clause 5, wherein at least a part of the various circumferential stiffening spars is bonded, by means of a thermal process, to at least a part of the inner surface of the metallic skin.
    [0068] Clause 8. The aircraft according to clause 1, wherein the various circumferential stiffeners comprise corrugated stiffeners having a continuous, corrugated design.
    [0069] Clause 9. The aircraft according to clause 8, wherein the continuous, corrugated model comprises a series of contiguous and sequential peaks and valleys.
    [0070] Clause 10. The aircraft according to clause 1, wherein the plurality of circumferential reinforcing spars comprise a plurality of circumferential blade reinforcing spars.
    [0071] Clause 11. The aircraft according to clause 10, wherein the circumferential blade reinforcement spars comprise a top end cap and a web section.
    [0072] Clause 12. The aircraft according to clause 11, wherein the circumferential blade reinforcement spars are in the form of a capital letter T.
    [0073] Clause 13. The aircraft according to clause 1, wherein the various circumferential stiffening spars are attached to a keel beam or a fuselage crown beam.
    [0074] Clause 14. An aircraft, comprising: a fuselage comprising: a plurality of circumferential stiffening spars coupled to an inner surface of a metal fuselage skin and disposed substantially normal to a first axis; the metal fuselage skin extending longitudinally to the along the first axis; and at least one beam disposed generally parallel to the first axis; and a wing comprising: a metallic wing skin having an inner surface, a first end adjacent to the fuselage, and a second end distant from the first end, extending from the first end to the second end along a second generally linear axis; and a plurality of wing rib spars coupled to the inner surface of the metallic wing skin and disposed generally parallel to the second axis, wherein the plurality of wing rib spars substantially align with the plurality of circumferential rib spars.
    [0075] Clause 15. The aircraft according to clause 14, wherein the beam is a keel beam or a crown beam.
    [0076] Clause 16. The aircraft according to clause 14, wherein at least a part of the metallic coating is attached to the beam.
    [0077] Clause 17. The aircraft according to clause 14, wherein the metallic coating comprises a composite material.
    [0078] Clause 18. The aircraft according to clause 14, wherein the composite material comprises a carbon fiber reinforced composite or a carbon fiber reinforced thermoplastic composite.
    [0079] Clause 19. The aircraft according to clause 14, wherein at least a part of the various circumferential stiffening spars is bonded, by means of a thermal process, to at least a part of the inner surface of the metallic skin.
    [0080] Clause 20. The aircraft according to clause 14, wherein the various circumferential stiffeners comprise corrugated stiffeners having a continuous, corrugated pattern.
    [0081] Clause 21. The aircraft according to clause 20, wherein the continuous, corrugated design comprises a series of contiguous and sequential peaks and valleys.
    [0082] Clause 22. The aircraft according to clause 14, wherein the plurality of circumferential reinforcing spars comprise a plurality of circumferential blade reinforcing spars.
    [0083] Clause 23. The aircraft according to clause 22, wherein the circumferential blade reinforcement spars comprise a top end cap and a web section.
    [0084] Clause 24. The aircraft according to clause 23, wherein the circumferential blade reinforcement spars are in the form of a capital letter T.
    [0085] Clause 25. The aircraft according to clause 14, wherein at least a part of the various wing reinforcement spars is attached to the beam.
    [0086] Clause 26. The aircraft according to clause 14, wherein at least a part of the plurality of wing reinforcement spars is attached to at least a part of the plurality of wing reinforcement spars of a second wing.
    [0087] Clause 27. The aircraft according to clause 14, wherein the wing comprises a front spar and a rear spar.
    [0088] Clause 28. The aircraft according to clause 27, wherein the front spar and the rear spar are attached to the fuselage.
    [0089] Clause 29. The aircraft according to clause 28, wherein the forward spar and aft spar are attached to the fuselage at an angle to provide a sloping wing configuration.
    [0090] Clause 30. The aircraft according to clause 14, wherein the wing further comprises at least one rib coupled to at least a part of the wing reinforcement spars.
    [0091] Clause 31. A method of reinforcing an aircraft fuselage, the process comprising: arranging a plurality of circumferential reinforcement spars along a longitudinal axis of the aircraft fuselage such that the plurality of circumferential spars are positioned substantially parallel to each other si and substantially normal to the longitudinal axis; securing at least a portion of the circumferential stiffening stringers to a beam; and attach a metal fuselage skin to the various circumferential stiffening spars.
    [0092] Clause 32. The process according to clause 31, wherein the various wing reinforcement spars are attached to the fuselage of the aircraft.
    [0093] Clause 33. The process according to clause 31, wherein at least a portion of the fuselage metal skin is attached to the beam.
    [0094] Clause 34. The process according to clause 31, wherein the beam is a keel beam or a crown beam.
    [0095] Clause 35. The process according to clause 31, wherein the metallic fuselage skin comprises a composite material.
    [0096] Clause 36. The process according to clause 35, wherein the composite material comprises a carbon fiber reinforced composite or a carbon fiber reinforced thermoplastic composite.
    [0097] Clause 37. The process according to clause 31, wherein the coupling of at least a part of the fuselage metal skin to at least a part of the plurality of circumferential reinforcement spars comprises thermally bonding at least a part of the metal skin to fuselage to at least a portion of the various circumferential stiffening spars.
    [0098] Clause 38. The process according to clause 31, wherein the various circumferential gussets comprise corrugated gussets having a continuous, corrugated pattern.
    [0099] Clause 39. The process according to clause 38, wherein the continuous, corrugated pattern comprises a series of contiguous and sequential peaks and valleys.
    [00100] Clause 40. The process according to clause 31, wherein the plurality of circumferential reinforcing stringers comprise a plurality of circumferential blade reinforcing stringers.
    [00101] Clause 41. The process according to clause 40, wherein the circumferential blade reinforcement stringers comprise a top end cap and a web section.
    [00102] Clause 42. The process according to clause 41, wherein the circumferential blade reinforcement stringers are in the form of a capital letter T.
    [00103] Clause 43. The process according to clause 31, further comprising coupling a front spar and a rear spar of a wing to the fuselage of the aircraft.
    [00104] Clause 44. The process according to clause 43, further comprising attaching the front spar or aft spar to the fuselage of the aircraft at an angle to the longitudinal axis to provide a sloping wing configuration.
    [00105] Clause 45. A wing comprising: a spar comprising an elliptical aperture, sized and shaped to encompass an aircraft; and a metallic wing skin having an inner surface coupled to the spar.
    [00106] Clause 46. The wing according to clause 45, further comprising: a number of wing reinforcement spars attached to the inner surface of the metallic wing skin.
    [00107] Clause 47. The wing according to clause 46, wherein the various wing stiffening spars substantially align with the various circumferential stiffening spars of a fuselage section of the aircraft.
    [00108] Clause 48. The wing according to clause 45, wherein the spar comprises a front spar and a rear spar.
    [00109] Clause 49. The wing according to clause 48, wherein the front spar of the wing is coupled to a rear spar of a second wing.
    [00110] Clause 50. The wing according to clause 49, wherein the rear spar of the wing is coupled to a front spar of the second wing.
    [00111] Clause 51. The wing according to clause 50, wherein a third spar is attached to a rib of the wing and to a rib of the second wing.
    [00112] Clause 52. The wing according to clause 45, wherein the spar is configured for attachment to an aircraft fuselage, at an angle, to provide a sloping wing configuration.
    [00113] Clause 53. The wing according to clause 45, wherein the foci of the elliptical aperture are modified depending on an angular displacement between an aircraft fuselage and the wing.
    [00114] Clause 54. The wing according to clause 45, wherein the wing further comprises at least one rib.
    [00115] Based on what was presented above, it should be considered that the technologies for reinforcement of various components of an aircraft, using vertically oriented circumferential reinforcement spars, were presented in this descriptive report. The subject matter described above is provided by way of illustration only, and should not be construed as limiting. Various modifications and variations may be made to the subject matter described in the present specification, without following the exemplary embodiments and applications illustrated and described, and without departing from the true spirit and scope of the present invention, which is set forth in the claims that follow.
    权利要求:
    Claims (20)
    [0001]
    1. An aircraft (100), comprising: a plurality of circumferential corrugated reinforcement spars (412) having a continuous, corrugated pattern, disposed within a fuselage (402) of the aircraft (100) and configured substantially normal to a longitudinal axis of the aircraft. (100), wherein each circumferential corrugated reinforcement spar (412) comprises a top section (414a, 414b) and a bottom section (418a, 418b); and a metal cladding (404) coupled to the top section (414a, 414b) of each of the plurality of circumferential stiffening spars, characterized in that each spar top section (414a, 414b) is contiguous with a bottom section (414a, 414b). 418a, 418b) of an adjacent circumferential reinforcement spar.
    [0002]
    2. Aircraft (100) according to claim 1, characterized in that it further comprises at least one beam arranged generally parallel to the longitudinal axis, wherein the beam is a keel beam (110) or a crown beam (210). ).
    [0003]
    3. Aircraft (100) according to claim 2, characterized in that at least a part of the metallic coating (404) is attached to the beam (110, 210).
    [0004]
    4. Aircraft (100) according to any one of claims 1 to 3, characterized in that the metallic coating (404) comprises a composite material.
    [0005]
    5. Aircraft (100) according to any one of claims 1 to 4, characterized in that at least a part of the plurality of circumferential reinforcement spars (412) is connected, by means of a thermal process, to at least one part of the inner surface of the metallic coating (404).
    [0006]
    6. An aircraft (100) according to any one of claims 1 to 5, characterized in that the plurality of circumferential stiffening spars (412) are attached to a keel beam (110) or a crown beam (210) of the fuselage (402).
    [0007]
    7. An aircraft (100) according to claim 1, characterized in that the plurality of circumferential reinforcement spars (412) are coupled to an inner surface of the fuselage metal skin (404) and the fuselage metal skin (404) ) extending longitudinally along the first axis; the aircraft further comprising: at least one beam (110, 210) disposed generally parallel to the longitudinal axis; and a wing (108a) comprising: a metallic wing skin having an inner surface, a first end adjacent to the fuselage (402), and a second end distal to the first end, extending from the first end to the second end along a second, generally linear axis; and a plurality of wing stiffener spars (114) coupled to the inner surface of the metallic wing skin and disposed generally parallel to the second axis, wherein the plurality of wing stiffener spars (114) align substantially with the plurality of wing spars (114) substantially align with the plurality of wing spars. circumferential reinforcement (412).
    [0008]
    An aircraft (100) according to claim 7, characterized in that at least a part of the plurality of wing reinforcement spars (114) is attached to at least a part of a plurality of wing reinforcement spars (114) 116) of a second wing (108b).
    [0009]
    9. Aircraft (100) according to claim 7 or 8, characterized in that the wing (108) comprises a front spar (610) and a rear spar (612) attached to the fuselage (402).
    [0010]
    10. An aircraft (100) according to claim 9, characterized in that the front spar (610) or the rear spar (612) are attached to the fuselage (402) at an angle to provide an arrow wing configuration.
    [0011]
    11. Aircraft (100) according to any one of claims 7 to 10, characterized in that the wing (108) further comprises at least one rib (622) coupled to at least a part of the wing reinforcement spars (114) , 116).
    [0012]
    12. A method of reinforcing an aircraft fuselage (102, 402, 502, 700), the method comprising: arranging a plurality of circumferential corrugated reinforcing spars (412) along a longitudinal axis of the aircraft fuselage (402) , such that the plurality of circumferential corrugated gussets (412) are positioned substantially parallel to each other and substantially normal to the longitudinal axis, wherein each circumferential corrugated gusset (412) comprises a top section (414a, 414b) and a bottom section (418a, 418b); securing at least a portion of the circumferential stiffening stringers (412) to a beam (110, 210); and attaching a metal fuselage skin (404) to the top section (414a, 414b) of each of the plurality of circumferential stiffening spars (412), characterized in that each spar top section (414a, 414b) is contiguous with a bottom section (418a, 418b) of an adjacent circumferential reinforcement spar.
    [0013]
    Process according to claim 12, characterized in that it further comprises attaching a plurality of wing reinforcement spars (114, 116) to the fuselage of the aircraft (402).
    [0014]
    14. Process according to claim 12 or 13, characterized in that it further comprises attaching at least a part of the fuselage metal cladding (404) to the beam (110, 210).
    [0015]
    Process according to any one of claims 12 to 14, characterized in that the coupling of at least a part of the fuselage metallic skin (404) to at least a part of the plurality of circumferential reinforcement spars (412) comprises thermal bonding of at least a portion of the metal fuselage skin (404) to at least a portion of the plurality of circumferential stiffening spars (412).
    [0016]
    16. Process according to any one of claims 12 to 15, characterized in that it further comprises coupling a front spar (610) and a rear spar (612) of a wing (108) on the aircraft fuselage (402) .
    [0017]
    17. Process according to claim 16, characterized in that it further comprises fixing the front spar (610) or the rear spar (612) at an angle with respect to the longitudinal axis, to provide an arrow wing configuration.
    [0018]
    18. Process according to claim 16, characterized in that at least one of the front spar (610) or the rear spar (612) comprises an elliptical opening (614, 616), sized and formed to encompass the fuselage of the aircraft (402); and a metallic wing skin having an inner surface coupled to the spars (610, 612).
    [0019]
    A method as claimed in claim 13, characterized in that the plurality of wing stiffener spars (114, 116) substantially align with the plurality of circumferential stiffener spars (412) of the fuselage section (402).
    [0020]
    20. Process according to claim 18, characterized in that the foci of the elliptical aperture are modified, depending on an angular displacement between an aircraft fuselage and the wing.
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    同族专利:
    公开号 | 公开日
    EP2772427A1|2014-09-03|
    US20140145031A1|2014-05-29|
    AU2013251284B2|2017-06-15|
    JP6628955B2|2020-01-15|
    CA2831571A1|2014-05-26|
    KR20140067901A|2014-06-05|
    US9145197B2|2015-09-29|
    CA2831571C|2016-07-05|
    AU2013251284A1|2014-06-12|
    EP2772427B1|2017-05-10|
    RU2649839C2|2018-04-04|
    CN103832575A|2014-06-04|
    BR102013030182A2|2014-10-07|
    RU2013149494A|2015-05-20|
    JP2014111437A|2014-06-19|
    KR102066754B1|2020-01-15|
    CN103832575B|2017-07-28|
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    法律状态:
    2014-10-07| 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]|
    2021-11-16| B09A| Decision: intention to grant [chapter 9.1 patent gazette]|
    2022-01-18| 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 25/11/2013, OBSERVADAS AS CONDICOES LEGAIS. |
    优先权:
    申请号 | 申请日 | 专利标题
    US13/685,024|US9145197B2|2012-11-26|2012-11-26|Vertically integrated stringers|
    US13/685,024|2012-11-26|
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