Composite material wing and fuselage butt joint frame

Abstract

The invention belongs to the technical field of water surface aircraft composite materials, and relates to a composite material wing and fuselage butt joint frame. Comprising a metal connecting joint, a composite material n-shaped frame, a pressing joint sleeve and a composite material fuselage skin. The composite material n-shaped frame is connected with the fuselage skin in a glued mode, the metal connecting joint is connected with a web plate of the composite material n-shaped frame through a fastener, and lugs of the metal connecting joint, frame edge strips of the composite material n-shaped frame and the fuselage skin are connected in a glued mode through mucilage glue in a pairwise mode and connected into a whole through fasteners. And the composite fuselage skin is notched to realize connection of the all-composite fuselage and the wings. The fuselage skin of the water surface aircraft is of a foam sandwich structure, the fuselage skin needs water sealing, and the problems of corrosion protection and connection and sealing of fasteners can be better solved by adopting the n-shaped composite material frame.

Description

Technical Field

[0001] This invention belongs to the field of composite material technology for surface aircraft, and relates to a composite material wing-fuselage docking frame. Background Technology

[0002] The wing-fuselage joint frame of a seaplane typically uses a single machined metal frame structure. This metal frame structure is heavy, and fasteners are essential for its connection to the fuselage. Furthermore, the thermal stress at the connection between the metal and composite fuselage is severe, and the composite fastener connection significantly increases the structural weight and is also a stress concentration area, impacting the aircraft's lifespan. Particularly during takeoff, landing, and taxiing, the fuselage frame is subjected to severe impact forces, which can easily lead to serious accidents.

[0003] The wing-fuselage joint metal frame is a common structural form used in aircraft. The wing-fuselage joint area is subjected to large loads and complex stresses. Metal reinforcement frames have good design experience. However, more and more aircraft structures are now using composite material structures. Thermal stress and stress concentration at the connection between metal and composite materials have become prominent problems. At present, aircraft are increasingly inclined to adopt all-composite material structures.

[0004] The main drawbacks of the existing structure are: severe thermal stress at the connection between the metal frame and the composite fuselage; stress concentration and fatigue-sensitive areas at the connection between the composite material and the metal fasteners; and the increased weight of the aircraft structure due to the metal frame. Summary of the Invention

[0005] Purpose of the invention For metal machined frames, this invention patent, based on the characteristics of all-composite aircraft, adopts a composite material Z-shaped frame and metal connecting joints. The composite material Z-shaped frame is glued to the fuselage skin, and the metal connecting joints are fastened to the web plate of the composite material Z-shaped frame. The metal connecting joints are further reinforced with lugs and connected to the frame edge strip of the composite material Z-shaped frame and the fuselage skin through adhesive bonding and fastening. The composite material fuselage skin has notches to achieve the connection between the all-composite fuselage and wings.

[0006] Technical solution A composite material wing-fuselage docking frame for aircraft includes a metal connecting joint, a composite material Z-shaped frame, a crimp fitting sleeve, and a composite material fuselage skin. The composite material Z-shaped frame is adhesively bonded to the fuselage skin. The metal connecting joint is connected to the web of the composite material Z-shaped frame via fasteners. The lugs of the metal connecting joint are adhesively bonded to the frame edge strip of the composite material Z-shaped frame and the fuselage skin in pairs, and are then connected as a single unit using fasteners. The composite material fuselage skin has notches to achieve a fully composite material fuselage-wing connection.

[0007] Furthermore, the adhesive is cured at room temperature for 7 days after bonding.

[0008] Furthermore, there are two metal connecting joints, which are symmetrically arranged on the composite material Z-shaped frame along the central axis of the fuselage.

[0009] Furthermore, it also includes a crimp fitting sleeve, which is disposed on the metal connector for transmitting loads. The metal connector is made of aluminum alloy, and the crimp fitting sleeve is made of stainless steel. Furthermore, the composite material is specifically a carbon fiber fabric epoxy composite material.

[0010] Furthermore, the overall shape of the composite material Z-shaped frame is determined by the shape of the skin at the junction of the aircraft fuselage and wings.

[0011] Furthermore, the connecting area of ​​the Z-shaped frame joint is locally thickened to 7mm, the force diffusion area is 7mm to 5mm, and the remaining parts are 3mm. The thickness of the metal connecting joint is 6mm, and the thickness of the fuselage skin in the connecting area is 2.2mm.

[0012] Technical effect Compared to existing technologies, the advantages of this invention are as follows: The fuselage skin of this surface aircraft is a foam sandwich structure, requiring a watertight seal. Using a Z-shaped composite material frame better solves corrosion protection and avoids sealing issues related to fastener connections. The composite material Z-shaped frame is bonded to the fuselage skin, and the wing-fuselage connection is achieved by adding a wing-fuselage connection joint, solving the problem of concentrated force connection. The bonded connection between the Z-shaped composite material frame and the composite material skin solves the thermal stress problem of connecting traditional concentrated force metal frames to composite material fuselages, and also reduces structural weight. The Z-shaped frame is 80mm wide and 80mm high. Its dimensions are designed to meet the installation requirements of the wing-fuselage connection joint. The two flanges of the Z-shaped frame are each 40mm wide, ensuring strong bonding with the composite material fuselage skin and meeting the requirements for load transfer from the Z-shaped frame to the fuselage skin. The thickness of the Z-shaped frame in the concentrated force connection area of ​​the wing-fuselage joint is increased to 7mm, while the thickness of the concentrated force diffusion area transitions from 7mm to 5mm and then to 3mm to meet the requirements for concentrated force diffusion. The Z-shaped composite material frame uses T300 carbon fiber fabric prepreg material and is formed in a medium-temperature autoclave. The fuselage bulkhead experiences the most severe load under the step-by-step water immersion condition, with maximum tensile strain, maximum compressive strain, and maximum shear strain of 1647με, 1230με, and 2377με, respectively, meeting the strength requirements. The 7mm thickness of the concentrated force connection area between the Z-shaped frame and the wing-fuselage joint meets the fastener connection compressive strength requirements. Attached Figure Description

[0013] Figure 1 Layout diagram of the "Z" shaped frame; Figure 2 The U-shaped frame is connected to the wing and fuselage joints.

[0014] Among them: metal connecting joint (1), composite material Z-shaped frame (3), crimping sleeve (4), composite material fuselage skin (5). Detailed Implementation

[0015] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the embodiments of the present invention. Obviously, the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0016] The features and illustrative embodiments of various aspects of the present invention will now be described in detail. In the following detailed description, numerous specific details are set forth in order to provide a thorough understanding of the invention. However, it will be apparent to those skilled in the art that the invention may be practiced without requiring some of these specific details. The following description of embodiments is merely intended to provide a better understanding of the invention by illustrating examples of the invention. The invention is by no means limited to any specific setup and method set forth below, but covers any improvements, substitutions, and modifications to the structures, methods, and devices without departing from the spirit of the invention. In the following description, well-known structures and techniques are not shown to avoid unnecessarily obscuring the invention.

[0017] In the description of this invention, it should be noted that the directions or positional relationships indicated by terms such as "center," "upper," "lower," "left," "right," "vertical," "horizontal," "inner," and "outer" are based on the stated directions or positional relationships and are only for the convenience of describing and simplifying the invention, and should not be construed as limiting the invention. Furthermore, the use of ordinal numbers (e.g., "first and second," etc.) is for distinguishing objects and is not limited to this order, and should not be construed as indicating or implying relative importance.

[0018] In the description of this invention, it should be noted that, unless otherwise explicitly specified and limited, the terms "installation," "connection," and "linking" should be interpreted broadly, encompassing both direct connection and indirect connection via an intermediate medium. Those skilled in the art can understand the specific meaning of these terms in this invention based on the specific circumstances.

[0019] It should be noted that, unless otherwise specified, the embodiments of the present invention and the features thereof can be combined with each other, and the various embodiments can be referenced and cited in each other. The present invention will now be described in detail with reference to and in conjunction with the embodiments.

[0020] Example This invention relates to a composite material wing-fuselage docking frame for an aircraft, comprising a metal connecting joint (1), a composite material Z-shaped frame (3), a crimping sleeve (4), and a composite material fuselage skin (5). The composite material Z-shaped frame (3) is glued to the fuselage skin (5), and the metal connecting joint (1) is connected to the web of the composite material Z-shaped frame (3) by fasteners. The lugs of the metal connecting joint (1) are glued to the frame edge strip of the composite material Z-shaped frame (3) and the fuselage skin (5) in pairs, and are connected as a whole by fasteners. The composite material fuselage skin (5) has a notch to achieve a full composite material fuselage-wing connection.

[0021] In one embodiment of the present invention, preferably, the adhesive is cured at room temperature for 7 days. In one embodiment of the present invention, preferably, there are two metal connecting joints (1), which are symmetrically arranged on the composite material Z-shaped frame (3) along the central axis of the fuselage.

[0022] In one embodiment of the present invention, preferably, a crimp fitting sleeve (4) is further included, which is disposed on the metal connector (1) for transmitting load. The metal connector is made of 7050 aluminum alloy, and the crimp fitting sleeve is made of 15-5PH stainless steel. In one embodiment of the present invention, preferably, the specific material of the composite material is a carbon fiber fabric epoxy composite material.

[0023] In one embodiment of the present invention, preferably, the overall shape of the composite material Z-shaped frame (3) is determined by the shape of the skin at the connection between the aircraft fuselage and the wing.

[0024] The dimensions of the U-shaped frame, the parameters of the web and flange are determined based on the strength requirements to determine the fastener specifications, and then the dimensions of the fastener connection area are determined based on the fastener nailing point requirements. Fasteners are present in the area where the Z-shaped frame connects to the metal connector.

[0025] The joint area of ​​the Z-shaped frame is locally thickened to 7mm, the force diffusion area is 7mm to 5mm, and the rest is 3mm. The thickness of the metal joint is 6mm, and the thickness of the fuselage skin in the connection area is 2.2mm.

[0026] The above detailed embodiments are a description of the present invention. It should not be considered that the specific embodiments of the present invention are limited to these descriptions. For those skilled in the art, several simple deductions and substitutions can be made without departing from the concept of the present invention, and all of these should be considered to fall within the protection scope of the present invention.

Claims

1. A composite material wing-fuselage docking frame for aircraft, characterized in that, It includes metal connecting joints, composite material Z-shaped frames, crimp fittings, and composite material fuselage skin; the composite material Z-shaped frames are bonded to the fuselage skin, the metal connecting joints are connected to the web of the composite material Z-shaped frames by fasteners, the lugs of the metal connecting joints are bonded to the frame edge strips of the composite material Z-shaped frames and the fuselage skin by adhesive, and are connected as a whole by fasteners, and the composite material fuselage skin has notches to achieve the connection of the entire composite material fuselage and wings.

2. The docking frame as described in claim 1, characterized in that, The adhesive is cured at room temperature for 7 days after bonding.

3. The docking frame as described in claim 1, characterized in that, The metal connecting joints are two in number, symmetrically arranged on the composite material Z-shaped frame along the central axis of the fuselage.

4. The docking frame as described in claim 1, characterized in that, It also includes a crimp fitting sleeve, which is disposed on the metal connection joint for transmitting load; the metal connection joint is made of aluminum alloy, and the crimp fitting sleeve is made of stainless steel.

5. The docking frame as described in claim 1, characterized in that, The composite material is specifically a carbon fiber fabric epoxy composite material.

6. The docking frame as described in claim 1, characterized in that, The overall shape of the composite material Z-shaped frame is determined by the shape of the skin at the junction of the aircraft fuselage and wings.

7. The docking frame as described in claim 1, characterized in that, The joint area of ​​the Z-shaped frame has a local thickness of 7mm, the force diffusion area has a thickness of 7mm-5mm, and the rest of the area has a thickness of 3mm.

8. The docking frame as described in claim 1, characterized in that, The metal connector is 6mm thick, and the fuselage skin in the connection area is 2.2mm thick.

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