An aircraft wing-body junction framework

Abstract

This application relates to the field of aircraft structural technology, and provides an aircraft wing-body connection frame, comprising multiple composite material profiles connected by an adhesive layer of predetermined thickness, the thickness of which is configured to compensate for assembly tolerances. The layup structure of the composite material profiles includes first-direction fibers for bearing axial loads and second-direction fibers for bearing shear and torsional loads; the thickness ratio of the first-direction fibers to the second-direction fibers is differentiated according to the stress conditions of the composite material profiles. This application achieves a balance between lightweight frame and high load-bearing capacity, improving assembly accuracy and mass production efficiency.

Description

Technical Field

[0001] This application relates to the field of aircraft structure technology, specifically to an aircraft wing-body connection frame. Background Technology

[0002] In aircraft, the wing-body coupling frame, as the main load-bearing component connecting the left and right wings and the fuselage, needs to effectively transfer the aerodynamic loads of the wings to the fuselage structure. In existing technologies, the main body of the wing-body coupling frame is typically made of metal materials such as aluminum alloy or steel, formed into a beam structure through machining, and then bolted together with longitudinal and transverse beams to form a grid-like frame. However, metal machining processes are time-consuming, and surface anti-corrosion treatment is cumbersome, resulting in long production cycles per unit and low assembly efficiency, making it difficult to meet the demands of mass production for capacity and low cost. Simultaneously, the high density of metal materials makes the frame heavy, creating a bottleneck for lightweighting; furthermore, its poor corrosion resistance in high-temperature, high-humidity, and high-salt-spray environments, coupled with advanced surface treatments, further limits the improvement of production efficiency and cost reduction. Therefore, there is an urgent need for a wing-body coupling frame structure that can balance lightweighting, high load-bearing capacity, corrosion resistance, and efficient mass production. Summary of the Invention

[0003] To address the problems of low production efficiency, heavy weight, and difficulty in balancing high-precision assembly and complex stress conditions in existing wing-body connection frames, this application proposes an aircraft wing-body connection frame and an aircraft, achieving lightweight, high load-bearing capacity, corrosion resistance, and efficient mass production and high-precision assembly of the frame.

[0004] To achieve the above objectives, the present invention adopts the following technical solution: A wing-body connection frame for an aircraft, characterized in that it comprises: Multiple composite material profiles are connected to each other by an adhesive layer of predetermined thickness; The composite material profile has a layup structure including first-direction fibers for bearing axial loads and second-direction fibers for bearing shear and torsional loads. The thickness ratio of the first direction fiber to the second direction fiber is configured differently according to the stress conditions of the composite material profile.

[0005] Furthermore, the composite material profile includes: A plurality of crossbeams, the crossbeams being arranged sequentially along the wingspan direction; A plurality of longitudinal beams, said longitudinal beams including those arranged sequentially along the course of travel; and Beam connectors; The crossbeams, longitudinal beams, and beam connectors are all of uniform cross-section profiles, and the crossbeams and longitudinal beams are cross-connected by beam connectors and adhesive layers.

[0006] Furthermore, the adhesive layer (300) has a gap for compensating for the accuracy of the structural connection surface, and the preset thickness is 0.5 mm to 3 mm.

[0007] Furthermore, the first directional fiber (100) is a 0° fiber, and the second directional fiber (200) is a ±45° fiber.

[0008] Furthermore, the thickness percentage of the second-direction fiber (200) in the crossbeam (101) is 10% to 60%, the thickness percentage of the second-direction fiber (200) in the longitudinal beam (201) is 10% to 60%, and the thickness percentage of the second-direction fiber (200) in the beam connector (601) is 50%. Furthermore, the thickness percentage of the first-direction fiber is 40% to 80% in the crossbeam assembly, 40% to 80% in the longitudinal beam assembly, and 20% to 60% in the beam connector.

[0009] Furthermore, the cross-sectional shape of the beam is selected from one of the following: square, sun, field, or I-beam. The cross-sectional shape of the longitudinal beam is selected from one of the following: square, sun, field, or I-beam. The cross-sectional shape of the beam connector is selected from one of the following: L-shaped, C-shaped, or trihedral box.

[0010] Furthermore, it also includes a local reinforcement component, wherein the composite material profile has a web, and the local reinforcement component is mechanically connected to the web; The local reinforcement component is a metal T-shaped bushing, the surface of which is subjected to anodizing or electrophoretic anti-corrosion treatment.

[0011] Furthermore, it also includes a reinforcing plate assembly, which is fitted at the junction of the crossbeam and the longitudinal beam (201).

[0012] Furthermore, the reinforcing sheet assembly is a continuous fiber reinforced resin matrix composite laminate structure; the fiber layup angle of the reinforcing sheet assembly includes ±45°, 0° and 90°, and is configured as a quasi-isotropic layup structure, with the thickness ratio of each fiber direction being 20% ​​to 40%.

[0013] Beneficial effects: The wing-body connection frame for aircraft provided in this application replaces the traditional metal structure with composite material profiles. By utilizing the high specific strength and high specific stiffness characteristics of continuous fiber reinforced resin matrix composites, the structural weight is significantly reduced while ensuring load-bearing capacity, breaking through the lightweight bottleneck of metal materials. Moreover, the composite material itself has excellent corrosion resistance, eliminating the need for additional surface treatment processes, thereby improving production efficiency and reducing manufacturing costs.

[0014] By setting an adhesive layer with a preset thickness to connect the various composite material profiles, the thickness margin of the adhesive layer is used to compensate for the dimensional tolerances in the profile manufacturing and assembly process. This avoids the problems of frequent repairs and drilling during the assembly of traditional metal frames, and realizes high-precision and rapid assembly of the frame, which greatly improves assembly efficiency and product consistency.

[0015] By designing composite profiles as constant cross-section profiles, they are adapted to continuous automated molding processes such as pultrusion. Compared with traditional metal machining or autoclave processes, this significantly improves material utilization and production cycle time, meeting the needs of large-scale stable production. At the same time, the constant cross-section structure also provides a physical basis for the uniform application of adhesive layers, further ensuring the compensation effect of assembly accuracy.

[0016] By differentiating the thickness ratio of first-direction fibers to second-direction fibers in the ply structure, the designability of the anisotropy of composite materials is fully utilized, so that the fiber arrangement of beams, longitudinal beams and connectors can be precisely matched with their respective stress characteristics under actual working conditions (such as axial force, shear force, torque, etc.). This maximizes structural efficiency while avoiding material redundancy, and achieves a balance between lightweight and high load-bearing capacity.

[0017] By setting up local reinforcing components and reinforcing plates, and using quasi-isotropic lay-ups or specific fiber configurations at key connection nodes, the strength weakening problem of composite profiles in stress concentration areas such as openings and connections is effectively solved, ensuring the integrity and reliability of the overall skeleton structure. At the same time, the discontinuous or continuous structural design of the longitudinal beams provides flexible assembly solutions to adapt to different body integration requirements. Attached Figure Description

[0018] Figure 1 This is a schematic diagram of the structure of an embodiment of the present invention; Figure 2 This is an exploded view of an embodiment of the present invention; Figure 3 yes Figure 1 A cross-sectional view; Figure 4 This is a cross-sectional schematic diagram of the beam according to an embodiment of the present invention; Among them, 10 is the web; 20 is the reinforcing sheet assembly; 30 is the local reinforcing assembly; 100 is the first direction fiber; 101 is the crossbeam; 200 is the second direction fiber; 201 is the longitudinal beam; 300 is the adhesive layer; and 601 is the beam connector. Detailed Implementation

[0019] To make the objectives, technical solutions, and advantages of this application clearer, the technical solutions in the embodiments of this application will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of this application.

[0020] Unless otherwise defined, all technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this application pertains.

[0021] This embodiment provides an aircraft wing-body connection frame, which serves as the main load-bearing hub connecting the wing and fuselage, aiming to solve the technical problems of traditional metal frames, such as large weight, low assembly efficiency, and poor corrosion resistance. The aircraft wing-body connection frame includes multiple composite material profiles, which are connected by an adhesive layer 300 of a preset thickness. The thickness of the adhesive layer 300 is configured to compensate for assembly tolerances. Specifically, the composite material profiles, as the basic load-bearing units of the frame, are made of continuous fiber-reinforced resin-based composite materials. Utilizing their high specific strength and high specific stiffness characteristics, the structural weight is significantly reduced while meeting the same load-bearing requirements. Simultaneously, the composite material itself possesses excellent corrosion resistance, eliminating the need for cumbersome surface anti-corrosion treatments required for metal structures, thereby improving production efficiency and reducing life-cycle costs. The adhesive layer 300 here not only serves a bonding and fixing function but, more importantly, acts as a designable tolerance compensation interface. Because dimensional deviations are unavoidable during the manufacturing and assembly of large frames, the pre-thickness adhesive layer 300 can absorb these accumulated tolerances through its own thickness margin. This eliminates the need for traditional fitting or drilling when connecting various profiles, achieving high-precision and rapid assembly without the need for fitting, significantly improving assembly efficiency and product consistency. It should be understood that the pre-thickness of the adhesive layer 300 is not a fixed value, but rather a design that matches the manufacturing precision level and assembly clearance requirements of the profiles, as long as the tolerance compensation function is achieved.

[0022] The layup structure of composite profiles includes first-direction fibers 100 for bearing axial loads and second-direction fibers 200 for bearing shear and torsional loads. Specifically, the first-direction fibers 100 and second-direction fibers 200 are defined based on their primary mechanical functions within the structure, rather than being limited to specific geometric angles. The first-direction fibers 100 are primarily arranged along the length of the component or the main force transmission path, responsible for resisting tensile and compressive axial loads; the second-direction fibers 200 are arranged at an angle relative to the main force transmission path, primarily responsible for resisting interlaminar shear, torsion, and lateral loads. This function-oriented fiber classification allows the layup design to accurately respond to the actual stress requirements of the component, avoiding performance waste or bottleneck effects caused by single-direction layups.

[0023] The thickness ratio of the first-direction fiber 100 to the second-direction fiber 200 is configured differently according to the stress conditions of the composite profile. Specifically, the differentiated configuration means that for profiles in different positions and with different functions in the skeleton, the thickness ratio of the two fibers is set independently according to their respective load spectrum. For example, for components that mainly bear axial tensile and compressive stress caused by bending moment, the thickness ratio of the first-direction fiber 100 is increased to maximize axial stiffness; while for connection node areas that mainly bear complex shear force or torque, the thickness ratio of the second-direction fiber 200 is appropriately increased to improve shear and torsional resistance. This material design strategy of allocation as needed fully utilizes the designability of the anisotropy of composite materials, ensuring the overall structural safety margin while eliminating material redundancy in non-stress directions, and achieving ultimate lightweight design.

[0024] Through the above technical solution, this embodiment organically combines composite material profiles, a pre-thickness adhesive layer 300, and a differentiated fiber layup configuration, resulting in a significant synergistic effect. The composite material profiles provide the material basis for lightweight and corrosion resistance; the pre-thickness adhesive layer 300 solves the assembly precision problem faced by composite material structures in mass production, making efficient automated assembly possible; and the differentiated fiber layup configuration optimizes structural performance under the process constraints of mass production of profiles with uniform cross-sections. The combined effect of these three elements enables the wing-body connection frame to overcome the weight bottleneck of traditional metal structures while simultaneously meeting the requirements of high-precision assembly, high load-bearing capacity, and low-cost mass production. It is particularly suitable for new aircraft platforms such as eVTOL and UAVs, which are sensitive to production capacity and lightweighting.

[0025] In one embodiment, multiple composite material profiles include several crossbeams 101 arranged along the wingspan direction, several longitudinal beams 201 arranged along the heading direction, and multiple beam connectors 601. The crossbeams 101, longitudinal beams 201, and beam connectors 601 are all profiles with uniform cross-sections and webs 10. One crossbeam 101 and one longitudinal beam 201 are cross-connected via beam connectors 601 and adhesive layers 300. Specifically, combined with... Figure 2 The exploded and cross-sectional views of the skeleton shown indicate that there are two crossbeams 101, extending along the wingspan of the aircraft, serving as the main load-bearing components to withstand wing bending moments. There are three longitudinal beams 201, arranged along the aircraft's heading, used to transfer axial loads on the fuselage and maintain the longitudinal stability of the skeleton. Beam connectors 6 are located at the intersections of the crossbeams 101 and longitudinal beams 201, connecting them to form a stable grid-like skeleton structure. In this embodiment, the crossbeams 101, longitudinal beams 201, and beam connectors 6 are all designed as profiles with constant cross-sections, meaning they have constant cross-sectional geometric parameters along their length. This constant cross-section design is, on the one hand, a necessary structural feature for adapting to the pultrusion molding process, enabling continuous and efficient production of continuous fiber-reinforced resin matrix composites at speeds of 0.3 m / min to 1 m / min, with a material utilization rate exceeding 95%, significantly superior to traditional metal machining or autoclave processes. On the other hand, the constant cross-sectional profile ensures geometric consistency of the adhesive mating surfaces throughout their entire length, providing a physical basis for the uniform application of the pre-set thickness adhesive layer 300, allowing for stable implementation of tolerance compensation and avoiding problems such as uneven adhesive layer thickness or stress concentration caused by abrupt changes in the profile cross-section. It should be understood that although this embodiment uses a grid-shaped frame as an example, in other embodiments, the topology of the frame can be adjusted to a star-shaped, trapezoidal, or other forms according to the aerodynamic layout of the aircraft. As long as each component maintains the characteristics of a constant cross-section profile and is connected by the adhesive layer 300, the aforementioned synergistic effect of mass production and high-precision assembly can be achieved.

[0026] In one implementation, an adhesive layer 300 is provided between the beam connector 601, the crossbeam 101, and the longitudinal beam 201. The adhesive layer 300 has a gap for compensating for the accuracy of the structural connection surface, and its preset thickness is 0.5mm to 3mm. The beam connector 601, the crossbeam 101, and the longitudinal beam 201 adopt a composite connection form combining adhesive bonding and mechanical fastening. Specifically, at the connection interface between the beam connector 601 and the crossbeam 101 and the longitudinal beam 201, this thickness range is an optimized range determined based on a large number of process tests: if the adhesive layer thickness is less than 0.5mm, it is difficult to fully absorb the warping deformation and cumulative assembly errors generated by the composite material profile during pultrusion curing, resulting in poor fit of the connection surface or even internal stress; if the adhesive layer thickness is greater than 3mm, although the tolerance compensation capability is enhanced, the shear modulus of the adhesive layer itself is low, and an excessively thick adhesive layer will generate large shear deformation under load, reducing the stiffness of the connection node, and is prone to generating bubble defects during the curing process. Therefore, a range of 0.5mm to 3mm ensures both assembly accuracy compensation and the mechanical performance and process reliability of the connection structure. In terms of connection method, this embodiment adopts a composite connection strategy with adhesive bonding as the primary method and mechanical fastening as a secondary method. The adhesive layer 300 bears the main load transfer function and provides sealing and corrosion protection, making the stress distribution in the connection area more uniform and avoiding the stress concentration problem at the fastener hole edges in purely mechanical connections. The mechanical fasteners (such as blind rivets or high-strength bolts) serve as safety redundancy, mainly used to prevent brittle peeling failure of the adhesive layer under accidental impact or long-term fatigue load. They also play a temporary positioning and pressurizing role during assembly, ensuring that the adhesive layer maintains the preset thickness before curing. This composite connection method fully leverages the advantages of both adhesive bonding and mechanical connection, achieving high efficiency, high reliability, and long service life of the connection node.

[0027] As an embodiment, the cross-sectional shapes of the cross beam 101 and the longitudinal beam 201 are selected from one of a square shape, a double-square shape, a four-square shape, or an I-shape; the cross-sectional shape of the beam connecting member 601 is selected from one of an L-shape, a C-shape, or a three-sided corner box form. Specifically, the selection of the cross-sectional shape depends on the specific force-bearing requirements and spatial constraints of the component. In the preferred solution of this embodiment, the cross beam 101 adopts a closed square cross section, which has excellent torsional rigidity and biaxial bending rigidity, and can effectively bear the complex combined bending and torsional loads transmitted from the wing; the longitudinal beam 201 also adopts a square cross section to ensure the axial load-bearing stability along the heading direction. In other alternative embodiments, when the cross beam 101 mainly bears unidirectional bending moment and is extremely sensitive to weight, an I-shaped cross section can also be selected, which increases the cross-sectional moment of inertia by concentrating the material on the upper and lower flanges to achieve extreme lightweighting; when a double-square or four-square cross section is selected, it is usually to provide a larger shear-resistant area or a multi-chamber function integration space within a limited height space. For the beam connecting member 601, this embodiment preferably adopts an L-shaped cross section, whose two right-angle sides are respectively attached to the web 10 or the flange of the cross beam 101 and the longitudinal beam 201, with a simple structure and easy pultrusion molding; in scenarios where the node force is more complex or higher connection rigidity is required, the connecting member can also adopt a C-shaped cross section to provide three-sided wrapping support, or adopt a three-sided corner box form to realize smooth transition and dispersion of multidirectional loads. It should be understood that the above list of cross-sectional shapes is not an exhaustive limitation, and any equal-section profile form that can be molded by a pultrusion process and can cooperate with the adhesive layer 300 to realize the tolerance compensation function shall be regarded as an equivalent alternative solution of the present application and included in the protection scope.

[0028] As an embodiment, the first-direction fibers 100 are 0° fibers, and the second-direction fibers 200 are ±45° fibers; the plurality of composite profile members include the beam connecting member 601, the cross beam 101 and the longitudinal beam 201, the layup structures of the beam connecting member 601, the cross beam 101 and the longitudinal beam 201 are all composed of the first-direction fibers 100 and the second-direction fibers 200, and the thickness proportions of the first-direction fibers 100 and the second-direction fibers 200 are different from each other among the cross beam 101, the longitudinal beam 201 and the beam connecting member 601. Specifically, in combination with Figure 2The lower part of the ply diagram illustrates that this embodiment specifies the functional fiber orientation defined in Example 1 as standard composite material ply angles. The 0° fibers are laid parallel to the profile length, serving as the first direction fiber 100. These fibers primarily bear the tensile and compressive loads caused by wing bending moments or fuselage axial forces, and are a core element ensuring the axial stiffness and strength of the frame. The ±45° fibers are laid alternately at ±45° angles relative to the profile axis, serving as the second direction fiber 200. These fibers primarily resist interlaminar shear stress, torsional loads, and compressive stress at the fastener hole edges, while also maintaining the stability of the profile cross-sectional shape. This binary ply system composed of 0° and ±45° fibers satisfies the pultrusion process's requirements for continuous fiber orientation and covers the main stress modes of the wing-body connection frame in actual service environments. More importantly, although the crossbeams 101, longitudinal beams 201, and beam connectors 601 use the same material system, the thickness ratio of 0° fibers to ±45° fibers inside them is not uniformly set. Instead, it is customized according to the specific role of each component in the force transmission path of the skeleton. This differentiated configuration breaks away from the traditional design practice of uniformly adopting conservative plying in pursuit of safety margins. It achieves a precise match between material properties and component stress characteristics, thereby minimizing redundant weight in non-load-bearing directions while ensuring structural safety.

[0029] In a preferred embodiment, the thickness percentage of the second-direction fiber 200 is 10% to 60% in the crossbeam 101, 10% to 60% in the longitudinal beam 201, and 40% to 60% in the beam connector 601; the thickness percentage of the first-direction fiber 100 is 40% to 80% in the crossbeam 101, 40% to 80% in the longitudinal beam 201, and 20% to 60% in the beam connector 601. Specifically, the above numerical ranges are the optimal balance points derived from optimization under typical aircraft wing-body connection conditions. For the crossbeam 101, as the main load-bearing structure of the wing root bending moment, it bears extremely high axial tensile and compressive stresses, thus requiring the highest 0° fiber content to provide sufficient axial modulus and strength. The corresponding ±45° fiber percentage is controlled between 10% and 60% to meet basic shear and torsional resistance requirements. For longitudinal beam 201, its main function is to transmit the axial load of the fuselage and coordinate the deformation of the crossbeam. The stress state is mainly axial force, but slightly less than that of crossbeam 101. Therefore, the 0° fiber ratio is set to 40% to 80%, and the ±45° fiber ratio is adjusted to 10% to 60% accordingly. While ensuring axial load-bearing capacity, the shear resistance is slightly improved to adapt to the complex node force transmission. As for beam connector 601, it is located in the stress concentration area where crossbeam 101 and longitudinal beam 201 intersect. It bears complex shear force, torque and eccentric load from multiple directions. Fiber reinforcement in a single direction is difficult to cope with multiaxial stress state. Therefore, the ±45° fiber ratio is increased to 50%, so that the 0° and ±45° fibers are configured with equal thickness. This gives the connector approximately quasi-isotropic in-plane mechanical properties, effectively preventing early shear failure or delamination at the node. It should be understood that the specific percentage values ​​mentioned above are only preferred examples for the load spectrum of a particular aircraft model. In actual engineering applications, designers can make adaptive adjustments based on the above values ​​according to the specific load conditions, safety factor requirements, and process windows of different aircraft. For example, when the torque component borne by the crossbeam 101 increases significantly, its ±45° fiber ratio can be appropriately increased to more than 10%; when the beam connector 201 mainly bears unidirectional shear force, its fiber ratio can also be fine-tuned to further optimize efficiency.

[0030] In one implementation, the aircraft wing-body connection frame also includes a local reinforcing component 30, which is disposed at the opening in the web 10. Specifically, in conjunction with Figure 2The exploded view of the skeleton shown indicates that mounting holes for connecting the wings are provided on the web plates 10 on both sides of the crossbeam 101 and longitudinal beam 201. Due to stress concentration caused by fiber discontinuity at the openings in the composite laminate, and because the compressive strength at the hole edge is much lower than that of metallic materials, direct connection via fasteners is highly susceptible to delamination or compressive failure. Therefore, this embodiment provides local reinforcing components 30 at the aforementioned opening locations, namely, outer reinforcing components and inner reinforcing components for the crossbeams. These reinforcing components, as heterogeneous material reinforcements, specifically target the weak points of the composite structure, ensuring the structural integrity of the connection points between the skeleton and the wings. It should be understood that although this embodiment uses the wing connection holes on the web plate 10 of the crossbeam 101 as an example, in other embodiments, the same principle can be used to provide local reinforcing components 30 wherever through holes are opened in the composite profile and bear large concentrated loads, such as the connection between the longitudinal beam end and the fuselage, or equipment mounting points. Their application should not be limited to the web plate 10 of the crossbeam 101.

[0031] The local reinforcing component 30 is a T-shaped bushing made of metal. The surface of the T-shaped bushing is treated with anodizing or electrophoretic corrosion protection. The T-shaped bushing has a flange, which is connected to the composite material profile by adhesive bonding or riveting. Specifically, the T-shaped bushing is made of lightweight, high-strength metal materials such as aluminum alloy. Its geometric shape is T-shaped, including a cylindrical section inserted into the opening of the web 10 and a flange that fits against the surface of the web 10. Metal was chosen instead of all composite materials for this bushing because metal has isotropic mechanical properties and excellent compressive strength, which can effectively resist the radial compressive force applied by fasteners and make up for the inherent defects of insufficient compressive strength at the hole edge of composite materials. At the same time, considering that the service environment of aircraft is usually accompanied by corrosive factors such as high humidity and high salt spray, and that direct contact between the metal bushing and carbon fiber composite material is prone to galvanic corrosion, the surface of the T-shaped bushing must be insulated and protected. This embodiment preferably employs anodizing or electrophoresis processes to form a dense and stable insulating oxide film or coating on the metal surface. This not only blocks the electrochemical corrosion pathway but also improves surface wear resistance and adhesion to the adhesive layer. Regarding the connection method, the flange and the composite web 10 are not solely fastened mechanically; instead, a combination of adhesive bonding or riveting is preferred. The adhesive layer 300 here continues the design concept of a preset thickness from the previous embodiment, absorbing the matching error between the machining accuracy of the metal bushing and the forming accuracy of the composite web 10, ensuring effective contact between the flange face and the web 10 throughout the entire area. Compared to pure bolt connections, the introduction of the adhesive layer 300 transforms the concentrated point load transmitted by the fasteners into a distributed surface load along the flange contact surface, significantly reducing the peak stress at the hole edge; while the mechanical fasteners serve as a safety redundancy, preventing the adhesive layer from failing under extreme impact or fatigue conditions. This integrated design of "metal bearing + adhesive layer force transmission + corrosion protection and isolation" fundamentally solves the reliability problem of perforated connections in composite materials. Furthermore, although this embodiment uses an aluminum alloy T-type bushing as a preferred example, in practical applications, the material of the local reinforcing component 30 can also be replaced with titanium alloy, stainless steel, or surface-metallized engineering plastics, depending on the corrosion resistance level or weight requirements; its structural form is not limited to the standard T-type. Any irregular bushing structure with flange expansion force transmission function, as long as it can disperse the load on the hole edge and form a reliable connection with the web plate 10, should be regarded as an equivalent alternative of this application and included in the scope of protection.

[0032] In one implementation, the aircraft wing-body connection frame also includes a reinforcing plate assembly 20, which is a continuous fiber-reinforced resin-based composite laminate structure. The fiber layup angles of the reinforcing plate assembly 20 include ±45°, 0°, and 90°, and it is configured as a quasi-isotropic layup structure, with the thickness ratio in each fiber direction ranging from 20% to 40%. The reinforcing plate assembly 20 is formed using a prepreg molding process. Specifically, combined with... Figure 2As shown, the reinforcing plate assembly 20 includes an upper reinforcing plate and a lower reinforcing plate, which are respectively attached to the intersection of the upper and lower edge strips of the crossbeam 101 and the longitudinal beam 201. Unlike the directional layup with 0° and ±45° as the main orientation in the crossbeam 101, longitudinal beam 201 and beam connector 601 in the previous embodiment, the reinforcing plate assembly 20 adopts a quasi-isotropic layup design composed of three fiber directions: ±45°, 0° and 90°, and the fiber thickness ratio of each direction is strictly controlled at 25%. This special layup configuration is specially designed based on the stress complexity at the connection node: in the intersection area of ​​the crossbeam and longitudinal beam, the load transmission path turns and redistributes, and the stress state exhibits significant multiaxiality and uncertainty. If the directional layup of the main beam is used, the node is prone to premature failure due to overload of fibers in one direction or underutilization of fibers in another direction. The quasi-isotropic layup, by balancing the fiber content in the three main directions, endows the reinforcing sheet with approximately equal stiffness and strength characteristics in all in-plane directions. This allows it to withstand multi-directional coupled loads from wing bending moments, fuselage axial forces, and assembly deviations without discrimination, effectively avoiding stress concentration in a single fiber direction and thus ensuring the continuity and reliability of the overall force transmission path of the skeleton. It should be understood that although this embodiment provides a preferred parameter of 25% thickness in each fiber direction, in actual engineering applications, this ratio can be finely adjusted within a certain range (e.g., 20% to 40%). As long as the overall layup maintains quasi-isotropic characteristics and can adapt to the multi-directional force requirements of the nodes, it should be considered an equivalent implementation of this application.

[0033] In terms of manufacturing process, the reinforcing plate assembly 20 is formed using a prepreg molding process, with a single-piece production cycle controlled within 10 minutes. This process choice forms an efficient capacity match with the pultrusion process of the main beam: the pultrusion process is responsible for the continuous and rapid manufacturing of the main body of the profile with the same cross-section, while the molding process focuses on the preparation of high-precision, short-cycle nodal reinforcements. The parallel production of both not only ensures the optimal matching of material properties among the various components of the frame, but also eliminates the waiting time for the whole aircraft assembly caused by the bottleneck of single component production at the system level, providing a solid process foundation for the large-scale and stable supply of the wing-body connection frame. In addition, the prepreg molding process also has the advantages of controllable fiber volume fraction, low porosity, and high dimensional accuracy, ensuring that the reinforcing plate can obtain a stable and reliable connection interface quality when bonded or riveted to the main beam, further improving the fatigue life and environmental adaptability of the frame assembly.

[0034] In one implementation, the longitudinal beam 201 is an intermittent structure interrupted by the transverse beam 101; or, the longitudinal beam 201 is a continuous beam structure, with a passageway on the transverse beam 101 for the longitudinal beam 201 to pass through. Specifically, in conjunction with Figure 2 and Figure 3As shown, this application provides two parallel technical solutions for the assembly topology of longitudinal beams 201 and crossbeams 101 in the wing-body connecting frame, to adapt to the final assembly process requirements of different aircraft models. In the first discontinuous assembly scheme, the longitudinal beam 201 is physically cut off along the flight direction by the crossbeam 101, forming three independent longitudinal beams. The ends of each segment of the longitudinal beam 201 are connected to the side of the crossbeam 101 through beam connectors 6 and fixed using the adhesive-riveting composite connection method mentioned in the previous embodiment. The core advantage of this segmented design is that it realizes the modularity of the frame structure: each segment of the longitudinal beam 201 can be independently pultruded, cut, and surface-treated, reducing the difficulty of single-piece manufacturing and the need for large tooling; in the assembly process, the segmented structure allows for greater freedom of position adjustment, and with the tolerance compensation function of the adhesive layer 300, it further reduces assembly stress; in addition, if a segment of the longitudinal beam is damaged during service, only the corresponding segment needs to be replaced without scrapping the entire frame, significantly improving maintenance economy. The cost is an increase in the number of nodes, which places higher demands on the precision of the connectors and the quality of the adhesive bonding.

[0035] In the second continuous assembly scheme, the longitudinal beam 201 adopts an integrated beam structure that runs through the transverse beam 101. In this case, the transverse beam 101 has a pre-designed clearance path matching the outer contour of the longitudinal beam 201 at its intersection with the longitudinal beam 201, allowing the longitudinal beam 201 to pass continuously through the transverse beam 101 without being interrupted. The longitudinal beam 201 and the transverse beam 101 can still be directly bonded together using the adhesive layer 300, or reinforced in the intersection area with local reinforcing sheet components 20 or beam connectors 601. Compared to the intermittent scheme, the continuous structure eliminates splicing nodes in the direction of the longitudinal beam 201, ensuring the absolute continuity of the axial force transmission path and avoiding stress concentration caused by abrupt changes in joint stiffness, thereby improving the overall longitudinal bending stiffness and fatigue life of the frame. Simultaneously, the reduction in the number of nodes simplifies the assembly process and shortens the final assembly cycle. However, this scheme requires that the crossbeam be precisely pre-reserved for clearance during molding, and the dimensional accuracy of the clearance must match the cross-sectional height of the longitudinal beam. This places more stringent process requirements on the pultrusion die design and curing deformation control of the crossbeam. It should be understood that the two assembly topologies mentioned above are not mutually exclusive, but rather adaptive choices based on the specific model's load level, production capacity planning, and maintenance strategy. Both fall within the scope of this application. For example, for verification models that prioritize extreme structural efficiency and have small production runs, a continuous approach can be preferred; while for commercial eVTOL or UAV platforms aimed at large-scale mass production and emphasizing low-cost maintenance, an intermittent approach offers greater overall advantages.

[0036] Regardless of which assembly scheme is adopted, the end area of ​​the longitudinal beam 201 is provided with openings for connecting to the machine body and fastener installation passages. Specifically, as shown in the figure... Figure 3As shown, the longitudinal beam 201 serves as the sole mechanical interface for docking with the forward and aft fuselage of the aircraft. Several sets of mounting holes are precision-machined on the web plates 10 or flanges in these end regions to accommodate fasteners such as high-strength bolts or titanium alloy screws, securely anchoring the frame to the fuselage section. The design of the fastener mounting pathways fully considers operational accessibility, ensuring that tightening operations can be completed smoothly even within the confined internal space of the fuselage. Simultaneously, the periphery of the end openings is typically locally thickened or reinforced according to stress distribution to resist concentrated reaction forces transmitted from the fuselage. This universal feature ensures that regardless of the transverse and longitudinal beam connection topology used internally, its interface standard with the external fuselage structure remains consistent, thereby improving the compatibility and interchangeability of the frame product with different fuselage configurations.

[0037] This embodiment also provides an aircraft, including a fuselage, wings, and a wing-body connecting frame as described in any of the foregoing embodiments, the wing-body connecting frame connecting the fuselage and the wings. Specifically, in combination with Figure 3 As shown, the wing-body connecting frame serves as the central load-bearing hub, installed at the junction of the forward and aft fuselages of the aircraft, with its top and sides respectively connecting to the left and right wings. During flight, the aerodynamic lift, drag, and bending moment loads generated by the wings are first transmitted to the crossbeam 101 of the frame, then through the beam connector 6 and adhesive layer 300 to the longitudinal beam 201, and finally distributed to the aft and rear fuselage structures through the mounting holes at the ends of the longitudinal beam 201. In this force transmission path, the pre-thickness adhesive layer 300 detailed in the aforementioned embodiment not only ensures assembly accuracy but also avoids stress concentration at the nodes through uniform stress distribution; while the differentiated fiber layup ensures that each component has optimal stiffness-strength matching in its main load-bearing direction, making the entire force transmission chain efficient and reliable. It should be understood that, although Figure 3 A specific fuselage docking configuration is shown, but in other embodiments, the frame can also be applied to aircraft with blended wing-body, tandem layout, or other configurations, as long as it serves to connect the wings and fuselage, it falls within the protection scope of this application.

[0038] To verify the effectiveness of the above technical solution, a static test was conducted on the composite material wing-body connection frame provided in this embodiment. Combined with... Figure 4The static load test device shown has its frame bolted to a central fixed bracket to simulate the fuselage connection interface. Lever arms (angle steel) are installed at the wing connection points on both the front and rear sides of the frame, and sandbags are suspended at the ends of the lever arms to simulate wing aerodynamic loads. Test results show that the maximum load-bearing capacity of the front beam node reaches 48,500 N, and the maximum load-bearing capacity of the rear beam node reaches 4,000 N, fully meeting the design requirements. More importantly, while achieving the aforementioned high load-bearing performance, the total weight of this composite material frame is only 4.0 kg, achieving a 23% weight reduction compared to the traditional aluminum alloy metal frame solution (5.2 kg). This weight reduction is not solely due to material replacement, but rather the synergistic effect of multiple factors, including the efficient material utilization of the uniform cross-section profile, the reduction of redundant weight through differentiated layup, and the optimization of the number of connectors through adhesive bonding. Furthermore, thanks to the automation of the pultrusion and molding processes, the production efficiency of this frame is several times higher than that of traditional metal machining solutions, and the overall cost of mass production is reduced by 23%, significantly enhancing the product's market competitiveness.

[0039] Based on the aforementioned superior performance indicators and manufacturing economics, the wing-body connection frame provided in this embodiment is particularly suitable for aircraft platforms that are highly sensitive to production capacity, lightweighting, and cost, such as electric vertical takeoff and landing (eVTOL) aircraft, industrial drones, and general aviation aircraft. In these application scenarios, aircraft often face the demands of high-frequency takeoffs and landings, complex load profiles, and large-scale commercial operations. Traditional metal frames struggle to balance performance and cost, while autoclave composite structures are limited by production efficiency. This application successfully breaks through this dilemma through integrated innovation in structure, process, and materials, providing a practical and feasible main load-bearing structure solution for the mass production of new aircraft. It should be understood that the above experimental data and application scenarios are merely illustrative examples intended to demonstrate the feasibility and superiority of the technical solution presented in this application, and do not constitute a limitation on the scope of protection of this application.

[0040] The above description is merely a specific embodiment of the present invention, but the scope of protection of the present invention is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in the present invention should be included within the scope of protection of the present invention. Therefore, the scope of protection of the present invention should be determined by the scope of the claims.

Claims

1. A wing-body connection frame for an aircraft, characterized in that, include: Multiple composite material profiles are connected to each other by an adhesive layer (300) of a predetermined thickness; The composite material profile has a layup structure including first direction fibers (100) for bearing axial loads and second direction fibers (200) for bearing shear and torsional loads. The thickness ratio of the first directional fiber (100) to the second directional fiber (200) is configured differently according to the stress conditions of the composite material profile.

2. The aircraft wing-body connection frame according to claim 1, characterized in that, The composite material profile includes: A plurality of crossbeams (101), the crossbeams (101) being arranged sequentially along the wingspan direction; A plurality of longitudinal beams (201), said longitudinal beams (201) comprising those arranged sequentially along the course of travel; and Beam connector (601); The crossbeam (101), the longitudinal beam (20), and the beam connector (601) are all of equal cross-section profiles. The crossbeam (101) and the longitudinal beam (201) are cross-connected by the beam connector (601) and the adhesive layer (300).

3. The wing-body connection frame of an aircraft according to claim 1, characterized in that, The adhesive layer (300) has a gap for compensating for the accuracy of the structural connection surface, and the preset thickness is 0.5 mm to 3 mm.

4. The wing-body connection frame of an aircraft according to claim 1, characterized in that, The first directional fiber (100) is a 0° fiber, and the second directional fiber (200) is a ±45° fiber.

5. The wing-body connection frame of an aircraft according to claim 2, characterized in that, The thickness percentage of the second-direction fiber (200) in the crossbeam (101) is 10% to 60%, the thickness percentage of the second-direction fiber (200) in the longitudinal beam (201) is 10% to 60%, and the thickness percentage of the second-direction fiber (200) in the beam connector (601) is 50%. The thickness percentage of the first directional fiber (100) in the crossbeam assembly (10) is 40% to 80%, the thickness percentage of the first directional fiber (100) in the longitudinal beam assembly is 40% to 80%, and the thickness percentage of the first directional fiber (100) in the beam connector (601) is 20% to 60%.

6. The wing-body connection frame of an aircraft according to claim 5, characterized in that, The cross-sectional shape of the beam (101) is selected from one of the following: square, sun, field, or gong. The cross-sectional shape of the longitudinal beam (201) is selected from one of the following: square, sun, field, or I-beam. The cross-sectional shape of the beam connector (601) is selected from one of the following: L-shaped, C-shaped, or trihedral box.

7. The wing-body connection frame of an aircraft according to claim 1, characterized in that, It also includes a local reinforcement component (30), the composite profile having a web (10), the local reinforcement component (30) being mechanically connected to the web (10).

8. The wing-body connection frame of an aircraft according to claim 7, characterized in that, The local reinforcement component (30) is a metal T-shaped bushing, the surface of which is subjected to anodizing or electrophoretic anti-corrosion treatment.

9. The wing-body connection frame of an aircraft according to claim 1, characterized in that, It also includes a reinforcing plate assembly (20), which is fitted at the junction of the crossbeam (101) and the longitudinal beam (201).

10. The wing-body connection frame of an aircraft according to claim 9, characterized in that, The reinforcing sheet assembly (20) is a continuous fiber reinforced resin matrix composite laminate structure; the fiber layup angle of the reinforcing sheet assembly (20) includes ±45°, 0° and 90°, and is configured as a quasi-isotropic layup structure, with the thickness ratio of each fiber direction being 20% ​​to 40%.

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