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    Splice joints for composite aircraft fuselages and other structures
    11.
    发明授权
    Splice joints for composite aircraft fuselages and other structures 有权

    公开(公告)号：US10689086B2

    公开(公告)日：2020-06-23

    申请号：US15459810

    申请日：2017-03-15

    Applicant: The Boeing Company

    Inventor： Jeffrey F. Stulc ,  Wallace C. Chan ,  Brian C. Clapp ,  Neal G. Rolfes

    IPC: B64C1/06 ,  B64C1/12 ,  B64C1/14 ,  B64C1/00

    Abstract: Structures and methods for joining composite fuselage sections and other panel assemblies together are disclosed herein. In one embodiment, a shell structure configured in accordance with the present invention includes a first panel portion positioned adjacent to a second panel portion. The first panel portion can include a first stiffener attached to a first composite skin, and the second panel portion can include a second stiffener attached to a second composite skin. The shell structure can further include a fitting extending across a first edge region of the first panel portion and a second edge region of the second panel portion. A first end portion of the fitting can be attached to the first stiffener and the first composite skin, and a second end portion of the fitting can be attached to a second stiffener and a second composite skin, to join the first panel portion to the second panel portion.

    SPLICE JOINTS FOR COMPOSITE AIRCRAFT FUSELAGES AND OTHER STRUCTURES
    14.
    发明申请
    SPLICE JOINTS FOR COMPOSITE AIRCRAFT FUSELAGES AND OTHER STRUCTURES 有权
    组合式飞机熔接和其他结构的接头

    公开(公告)号：US20160251071A1

    公开(公告)日：2016-09-01

    申请号：US14525218

    申请日：2014-10-28

    Applicant: The Boeing Company

    Inventor： Jeffrey F. Stulc ,  Wallace C. Chan ,  Brian C. Clapp ,  Neal G. Rolfes

    IPC: B64C1/06 ,  B64C1/14 ,  B64C1/12

    CPC classification number: B64C1/069 ,  B64C1/068 ,  B64C1/12 ,  B64C1/1492 ,  B64C2001/0072 ,  Y02T50/43 ,  Y10T29/49616 ,  Y10T29/49622 ,  Y10T29/49826 ,  Y10T428/19 ,  Y10T428/24331 ,  Y10T428/24612 ,  Y10T428/24628 ,  Y10T428/24661

    Abstract: Structures and methods for joining composite fuselage sections and other panel assemblies together are disclosed herein. In one embodiment, a shell structure configured in accordance with the present invention includes a first panel portion positioned adjacent to a second panel portion. The first panel portion can include a first stiffener attached to a first composite skin, and the second panel portion can include a second stiffener attached to a second composite skin. The shell structure can further include a fitting extending across a first edge region of the first panel portion and a second edge region of the second panel portion. A first end portion of the fitting can be attached to the first stiffener and the first composite skin, and a second end portion of the fitting can be attached to a second stiffener and a second composite skin, to join the first panel portion to the second panel portion.

    Abstract translation: 本文公开了将复合机身部分和其它面板组件连接在一起的结构和方法。 在一个实施例中，根据本发明构造的壳结构包括邻近第二面板部分定位的第一面板部分。 第一面板部分可以包括附接到第一复合皮肤的第一加强件，并且第二面板部分可以包括附接到第二复合皮肤的第二加强件。 外壳结构还可以包括跨过第一面板部分的第一边缘区域延伸的配件和第二面板部分的第二边缘区域。 配件的第一端部可以附接到第一加强件和第一复合表皮，并且配件的第二端部可以附接到第二加强件和第二复合表皮，以将第一面板部分连接到第二加强件 面板部分。

    Hat stringer closeout fitting and method of making same
    15.
    发明授权
    Hat stringer closeout fitting and method of making same 有权
    帽子纵梁配合及其制作方法

    公开(公告)号：US09399510B2

    公开(公告)日：2016-07-26

    申请号：US14464098

    申请日：2014-08-20

    Applicant: The Boeing Company

    Inventor： Corey A. Rolfes ,  Jeffrey F. Stulc

    IPC: B64C3/00 ,  B64C3/18 ,  B64C3/34 ,  B64D37/00 ,  B64D37/04 ,  B64D37/10

    CPC classification number: B64C3/182 ,  B64C3/34 ,  B64D37/005 ,  B64D37/04 ,  B64D37/10

    Abstract: A closeout fitting for a hat stringer includes a cover that has a fitting surface. The fitting surface is substantially complementary to at least a portion of an outer surface of the hat stringer. The cover is configured to couple against the hat stringer outer surface. The cover also includes a first portion configured to extend across a gap defined in a cap portion of the hat stringer. The closeout fitting also includes an insert. The insert includes a first outer perimeter surface that is substantially complementary to at least a portion of an interior surface of a channel defined by the hat stringer. The insert is configured to couple against the channel interior surface proximate the hat stringer closeout portion. The insert and the cover are formed with a suitable stiffness to limit a deformation of the hat stringer proximate the closeout portion.

    Abstract translation: 用于帽子纵梁的封闭装置包括具有配合表面的盖。 配合表面与帽子纵梁的外表面的至少一部分基本互补。 盖子被配置成与帽子纵梁外表面联接。 盖子还包括构造成延伸穿过限定在帽子纵梁的帽部分中的间隙的第一部分。 封闭件还包括插入件。 该插入件包括第一外周表面，该第一外周表面与由帽子纵梁限定的通道的内表面的至少一部分基本互补。 插入件被配置为在靠近帽子纵梁关闭部分的情况下联接到通道内表面。 插入件和盖子形成有适当的刚度，以限制靠近封闭部分的帽子纵梁的变形。

    Composite access door
    16.
    发明授权
    Composite access door 有权
    复合门

    公开(公告)号：US09073618B2

    公开(公告)日：2015-07-07

    申请号：US13875548

    申请日：2013-05-02

    Applicant: The Boeing Company

    Inventor： Richard B. Tanner ,  Brian H. Korenaga ,  Jeffrey F. Stulc ,  Ian C. Burford

    IPC: B64C1/14 ,  E04C2/54 ,  E05D15/00

    CPC classification number: B64C1/1461 ,  B64C1/1446 ,  E04C2/54 ,  E05D15/00

    Abstract: A method of installing a door in an access opening in a skin of an aircraft. A composite inner door panel is positioned against an inside face of the skin. A composite outer door panel is positioned against an outside face of the skin. A plurality of fasteners is installed through fastener openings in the composite outer door panel. The plurality of fasteners is tightened to form a chamfered interface between the skin and the composite outer door panel.

    Abstract translation: 一种将门安装在飞行器皮肤的进入口中的方法。 复合内门板抵靠皮肤的内表面定位。 复合外门板抵靠皮肤的外表面定位。 多个紧固件通过复合外门板中的紧固件开口安装。 多个紧固件被紧固以在皮肤和复合外门面板之间形成倒角接口。

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