A composite material fuselage panel for an aircraft

Abstract

The present invention provides a composite material fuselage panel for an aircraft, comprising a skin panel, a plurality of annular reinforcing frames distributed along the heading direction of the aircraft on the inner surface of the skin panel, and a plurality of segmented stringers distributed along the circumferential direction of the aircraft between adjacent annular reinforcing frames. The inner surface of the skin panel has a circumferential local reinforcement portion with a predetermined thickness formed at the intersection of the segmented stringers and the annular reinforcing frames, and the end of the circumferential local reinforcement portion extending along the heading direction overlaps with the end of the segmented stringer near the circumferential local reinforcement portion.

Description

Technical Field

[0001] This invention relates to the field of aviation, mainly to the field of composite material structure design for aircraft, and particularly to a composite material fuselage panel for aircraft. Background Technology

[0002] In aircraft where composite materials are the primary material for fuselage panels, various connection methods and complex stress distributions exist in certain areas of the fuselage panels. Existing technologies mostly employ reinforcement methods by arranging large-sized circumferential reinforcing frames and azimuth reinforcing beams on the inner surface of the composite fuselage panels. However, these reinforcement methods can negatively impact the continuity of azimuth components, such as ordinary stringers, on the fuselage skin panels, making it impossible to achieve continuous transfer of stringer loads from the fuselage panels to the skin panels. Summary of the Invention

[0003] Therefore, the technical problem to be solved by the present invention is to overcome the problem that the existing composite fuselage panels have a large number of circumferential reinforcing frames or azimuth reinforcing beams, which makes it inconvenient to install ordinary stringers on their skin panels. The present invention proposes a new composite fuselage panel for aircraft.

[0004] The present invention solves the above-mentioned technical problems through the following technical solution:

[0005] Specifically, the present invention provides a composite material fuselage panel for an aircraft. The composite material fuselage panel includes a skin panel, a plurality of annular reinforcing frames distributed along the heading direction of the aircraft on the inner surface of the skin panel, and a plurality of segmented stringers distributed along the circumferential direction of the aircraft between adjacent annular reinforcing frames. The inner surface of the skin panel has a circumferential local reinforcement with a predetermined thickness formed at the intersection of the segmented stringers and the annular reinforcing frames. The end of the circumferential local reinforcement extending along the heading direction overlaps with the end of the segmented stringer near the circumferential local reinforcement.

[0006] The composite material fuselage panel for aircraft provided in this application combines a structural design of annular reinforcing frame and segmented stringers, and connects the segmented stringers to the skin panel with fasteners. This not only avoids interference from the annular reinforcing frame on the installation of the segmented stringers on the skin panel, thereby achieving an effective connection between the segmented stringers and the annular reinforcing frame on the skin panel of the composite material fuselage panel, but also comprehensively considers various requirements such as assembly, load transfer, stiffness, and sealing at the intersection of the segmented stringers and the annular reinforcing frame.

[0007] By designing a circumferential local reinforcement at the intersection of the segmented stringers and the annular reinforcing frame of the skin panel, and making the circumferential local reinforcement overlap with the end of the segmented stringers, the stiffness continuity of the circumferential reinforcement, the segmented stringers, and the skin panel in the intersection area can be guaranteed, avoiding the interruption of the directional load transmitted by the segmented stringers. This enables the continuous transmission of directional loads on the composite fuselage panel and also significantly improves the sealing performance of the composite fuselage panel.

[0008] According to one embodiment of the present invention, in the overlapping region of the circumferential local reinforcement and the end of the segmented stringer, the upper surface of the circumferential local reinforcement and the lower surface of the segmented stringer are in contact with each other. By designing the upper and lower surfaces of the circumferential local reinforcement and the segmented stringer to be in contact with each other, the rigidity of the composite fuselage panel can be ensured, so that the stress on the composite fuselage panel can be evenly distributed on the segmented stringer and the circumferential local reinforcement, and the continuity of directional load transfer and load transfer perpendicular to the directional load direction on the composite fuselage panel can be achieved.

[0009] According to one embodiment of the present invention, the shape of the composite fuselage panel at the region where it connects with the aircraft wing is designed to substantially match the shape of the region where the wing contacts it. Designing the shape of the composite fuselage panel to match the shape of the region where the wing contacts it, i.e., the shape of the upper surface of the wing, facilitates the connection between the composite fuselage panel and the aircraft wing, thereby enabling the assembly of the fuselage structure and the wing structure.

[0010] According to one embodiment of the present invention, the inner surfaces of multiple connecting regions pre-set inside the skin panel are designed to have the same thickness relative to the theoretical outer surface of the composite fuselage panel, and local thickness reinforcement or local thickness reduction portions are designed on one side of the theoretical outer surface according to the magnitude of local stress in local areas. By using a combination of internal and external layering in the assembly areas of the composite fuselage panel, the complexity of the inner surface of the composite fuselage panel in the assembly area is reduced, thereby reducing the possibility of shape measurement difficulties and gap compensation due to assembly positioning tolerances between the reinforcing frame, beam and panel.

[0011] According to one embodiment of the present invention, the inner surface is designed as a flat inner surface at a predetermined distance from the theoretical outer surface of the composite fuselage panel. By designing the inner surface of the composite fuselage panel, which has numerous connecting structures and complex connection relationships, as a flat inner surface, large variations in the thickness of the inner surface or inner profile of the composite fuselage panel can be avoided, thereby firmly connecting and fixing the connecting structures to the composite fuselage panel, thus ensuring effective connection between the composite fuselage panel and other structural components. Moreover, by maintaining a predetermined distance between the inner surface and the theoretical outer surface, the basic structural strength and rigidity of the skin panel can be guaranteed, preventing the skin panel itself from failing during use.

[0012] According to one embodiment of the present invention, the flat inner surface of the skin panel is designed as a plane or a curved surface. By designing the flat inner surface as a plane or a curved surface, it can be matched with the surface in contact with the connecting structure. For example, if the surface in contact with the connecting structure is curved, and the flat inner surface is a curved surface with the same curvature as the aforementioned curved surface, it facilitates the movement or rapid positioning of the connecting structure on the flat inner surface of the skin panel, and reduces the assembly difficulty of the connecting structure on the skin panel mating surface.

[0013] According to one embodiment of the present invention, a localized thickness reinforcement is designed on one side of the theoretical outer surface at a region of high local stress in the skin panel. By designing the required localized thickness reinforcement on one side of the theoretical outer surface, the resulting localized thickness reinforcement constitutes part of the formed skin panel (or the reinforcement or reinforcement area itself is located inside the skin panel). Therefore, it has advantages over a pre-defined reinforcing sheet bonded to the outer surface of the skin panel, such as more reliable manufacturing quality, lower cost, and better mechanical properties of the connection surface.

[0014] According to one embodiment of the present invention, the areas with high local stress include the areas where the skin panel is connected to the front and rear spars of the central wing of the aircraft, or the areas where multiple connection structures are predetermined to be provided on the inner surface of the skin panel.

[0015] According to one embodiment of the present invention, a localized thickness reduction section is designed on one side of the theoretical outer surface in a region of low local stress on the skin panel. By designing a thickness reduction section in a localized region of low or no stress, the weight of the composite material fuselage panel is reduced, thereby reducing the overall weight of the panel structure.

[0016] According to one embodiment of the present invention, the region with low local stress includes the region where the skin panel is connected to the middle portion of the central wing.

[0017] According to one embodiment of the present invention, composite material fuselage panels are configured for the wing-body connection area.

[0018] Based on common knowledge in the field, the above-mentioned preferred embodiments can be combined arbitrarily to obtain various preferred embodiments of the present invention.

[0019] The positive and progressive effects of the above-described embodiments of the present invention are as follows:

[0020] 1. For composite fuselage panels with annular reinforcing frames, segmented stringers are used to absorb and transfer yaw loads, avoiding interference between the annular reinforcing frame and the stringer assembly. The segmented stringers can be firmly fixed to the skin panels via mechanical connections. Simultaneously, at stringer segments with significant yaw loads, localized circumferential reinforcements of a certain thickness and width covering the ends of the segmented stringers are designed on the inner surface of the skin panels. This enables effective connections between the segmented stringers and between the segmented stringers and the skin panels and annular reinforcing frames, ensuring the stiffness continuity of the annular reinforcing frames, skin panels, and segmented stringers in the intersection area, thereby achieving the connection and transfer of yaw loads and complex load transfer between components or structures.

[0021] 2. By using the same thickness inner surface design in the assembly areas of a large number of connecting structures on the skin panel, the complexity of the inner surface of the composite material panel in the assembly area can be reduced. This reduces the possibility of shape measurement difficulties and gap compensation due to azimuth or circumferential assembly positioning tolerances, and reduces the assembly difficulty of the mating surfaces of the docking connection structure, reducing the number of compensation shims required. This effectively improves the integrity of the composite material fuselage panel structure, effectively reduces the assembly process difficulty, and improves the overall assembly speed.

[0022] 3. Moreover, by designing internal and external layer loss in the skin panel, that is, designing the inner surface of a local area to be of the same thickness, and designing a local thickness reinforcement or local thickness reduction part on one side of the outer surface of the corresponding local area, it is possible to ensure that the composite material fuselage panel itself meets the strength and stiffness requirements, while also effectively reducing the overall weight of the composite material fuselage panel structure. Attached Figure Description

[0023] Figure 1 A schematic diagram of a composite material fuselage panel according to a preferred embodiment of the present invention is shown.

[0024] Figure 2 schematically shown Figure 1 A cross-sectional view of the composite material fuselage panel along direction A.

[0025] Figure 3 A partial schematic diagram of a composite fuselage panel with numerous connections is shown.

[0026] Figure 4 schematically shown Figure 3A cross-sectional view of the composite material fuselage panel along direction B. Detailed Implementation

[0027] 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 in conjunction with the accompanying drawings showing multiple embodiments according to this application. It should be understood that all other embodiments obtained by those skilled in the art based on the embodiments described in this application without creative effort will fall within the scope of protection of this application.

[0028] Unless otherwise defined, all technical and scientific terms used in this application have the same meaning as commonly understood by one of ordinary skill in the art to which this application pertains; the terms "comprising," "having," etc., in the specification, claims, and foregoing description of the drawings are open-ended terms. Thus, "comprising" or "having" means, for example, one or more parts or portions, having one or more parts or portions, but is not limited to having only these one or more parts or portions.

[0029] In the aerospace field, composite materials are widely used in aircraft fuselage structures, especially as the primary material for fuselage panels. It is known that aircraft fuselage panels often require the installation of numerous complex connection structures, necessitating that composite fuselage panels incorporate intricate connection interfaces to meet operational requirements. Furthermore, they themselves must possess sufficient rigidity and strength to fulfill these requirements.

[0030] To ensure the strength of composite fuselage panels, annular metal reinforcing frames distributed along the aircraft's heading and aircraft stringers or heading reinforcing beams distributed along the aircraft's circumferential direction are often arranged on the skin panels of the composite fuselage panels simultaneously. This ensures the strength of the resulting composite fuselage panels, enabling them to withstand loads from other aircraft components or impacts from the flight environment.

[0031] However, the presence of numerous large-sized annular metal reinforcing frame structures on the composite fuselage skin panels, particularly in the most complex area of ​​the aircraft fuselage structure—the junction between the fuselage and the wing—can negatively impact the installation of azimuth components, including stringers, on the skin panels. This can prevent the stringers from being securely mounted, or create gaps between the stringers and the skin panels, leading to easy tearing or failure of the skin panels at these gaps during use. Furthermore, engineering experience shows that this method of intersecting the stringers with the annular metal reinforcing frames often results in stress concentration at the overlap points and can also cause defects such as discontinuous load transfer between the composite fuselage panels.

[0032] Therefore, this application provides a novel composite material fuselage panel for aircraft. For example... Figure 1 As shown, the composite material fuselage panel includes a skin panel 1, multiple annular reinforcing frames 2 distributed along the flight direction of the aircraft on the inner surface of the skin panel 1, and multiple segmented stringers 3 distributed along the circumferential direction of the aircraft between adjacent annular reinforcing frames 2. Wherein, as... Figure 2 As shown, a circumferential local reinforcement 11 with a predetermined thickness is formed on the inner surface of the skin panel 1 at the segment position of the segmented stringer 3 where the yaw load is large, that is, at the position where the segmented stringer 3 intersects with the annular reinforcing frame 2. The end of the circumferential local reinforcement 11 extending along the yaw direction overlaps with the end of the segmented stringer 3 near the circumferential local reinforcement 11.

[0033] like Figure 2 As shown, in the overlapping area at the ends of the circumferential local reinforcement 11 and the segmented stringer 3, the upper surface of the circumferential local reinforcement 11 and the lower surface of the segmented stringer 3 are in contact with each other to ensure the rigidity of the composite fuselage panel. This allows the stress on the composite fuselage panel to be evenly distributed on the segmented stringer 3 and the circumferential local reinforcement 11. The circumferential local reinforcement 11 then transfers the directional load to the adjacent segmented stringer 3 and the load perpendicular to the directional load direction to the annular reinforcing frame 2 and the skin panel 1. This achieves the continuity of directional load transfer and load transfer perpendicular to the directional load direction on the composite fuselage panel.

[0034] It should be understood that holes may be provided at corresponding positions of the annular reinforcing frame 2, the segmented stringer 3, and the skin panel 1, so that the annular reinforcing frame 2 and the segmented stringer 3 can be fixed to the corresponding positions of the skin panel 1 by fasteners.

[0035] In terms of external design, when composite fuselage panels are used to connect with the wing structure, such as Figure 3 As shown, the shape of the composite fuselage panel at the area where it connects with the aircraft's wing is designed to basically match the shape of the area where it contacts the wing. Through the aforementioned shape design, the shape fit and effective connection between the composite fuselage panel and the aircraft wing can be achieved, thereby realizing the assembly of the fuselage structure and the wing structure.

[0036] Meanwhile, the area where the fuselage structure connects to the wing structure is one of the most complex areas in an aircraft structure. Specifically, the connection relationships on the skin panel 1 of the composite fuselage panel are complex and the stress variations are large. Existing technologies typically employ methods such as bonding prefabricated reinforcing members to the corresponding side of the skin panel 1, or varying the thickness at the connection area of ​​the theoretical inner surface of the panel to achieve a localized reinforcement effect.

[0037] When thickness variations are designed for local reinforcement in numerous connection areas of the theoretical inner surface, the inner surface of the skin panel at the connection area will exhibit thickness variations along the circumferential direction. The connecting structure (e.g., circumferential reinforcement) that adheres to the skin panel in this area will experience circumferential displacement relative to its theoretical position during installation, resulting in significant gaps between the inner surface of the skin panel and the connecting structure in this area. To ensure proper fit and connection of the connecting structure to the skin panel, numerous shims with two parallelogram cross-sections are required to compensate for these gaps, increasing costs and failing to guarantee a stable connection between the connecting structure and the skin panel. More seriously, when thickness variations are designed for local reinforcement in the connection area of ​​the theoretical inner surface, the inner surface of the skin panel at the connection area will exhibit thickness variations along both the circumferential and yaw directions. In this case, when the circumferential reinforcement or connecting structure adheres to the skin panel in this area, it will experience circumferential and yaw offsets relative to its theoretical position. At this point, there will be gaps with complex shapes between the connecting structure installed on the skin panel and the skin panel. It is necessary to use shims with complex shapes, such as a combination of wedges and trapezoids, in different directions to compensate for the gaps. At this point, not only is it necessary to spend a lot of time and money to manufacture shims that meet the requirements to compensate for the gaps, but it is also difficult to effectively fill the gaps and achieve an accurate and effective connection between the skin panel and the connecting structure.

[0038] As mentioned earlier, when the circumferential and longitudinal reinforcing structures deviate in positioning, they can easily create complex gaps that are difficult to measure and effectively fill between the reinforcing structure and the wall panel. However, due to the complex distribution of internal forces or stresses in areas with complex connection relationships, it is difficult to implement the method of local reinforcement by bonding prefabricated reinforcing members on one side, and the quality is also difficult to guarantee. Making many thickness variations in the theoretical inner surface of the wall panel will make it difficult for the connection structure to fit the surface of the skin wall panel 1 for assembly and bonding to the composite material fuselage wall panel.

[0039] In response to the above problems, such as Figures 3 to 4 As shown, the composite material fuselage panel provided in this application also uses its theoretical outer surface Lo as a reference surface. The inner surfaces No at the predetermined connection areas inside the skin panel 1 are designed to have the same thickness. Furthermore, based on the magnitude of local stress in a local area, local thickness reinforcement or local thickness reduction is designed on one side of its theoretical outer surface Lo to form the actual outer surface Ls. Therefore, the inner surfaces No at the multiple connection areas predetermined inside the skin panel 1 are designed to have the same thickness relative to the theoretical outer surface Lo of the composite material fuselage panel, and the actual outer surfaces Ls at local areas with large or small local stress are designed to have local thickness reinforcement or local thickness reduction relative to their theoretical outer surface Lo.

[0040] Specifically, using the theoretical outer surface Lo of the composite material fuselage panel as a reference surface, the inner surface No is designed as a flat inner surface at a predetermined distance from the reference surface. Furthermore, the shape of this flat inner surface is designed to be either planar or curved, allowing it to match the shape of the surface of the connecting structure to which it is intended to be installed. For example, if the surface of the connecting structure is curved, and the flat inner surface here has the same curvature as the aforementioned curved surface, it facilitates the movement or rapid positioning of the connecting structure on the flat inner surface of the skin panel, and reduces the assembly difficulty of the connecting structure on the skin panel's mating surface.

[0041] Furthermore, a localized thickness reinforcement section is designed on the side of the theoretical outer surface Lo in areas of high local stress on the skin panel 1, that is, the part between the actual outer surface Ls and the theoretical outer surface Lo. For example... Figure 4 As shown, on the skin panel 1, in areas where complex connection structures such as the central wing joint 51 and the central wing No. 1 rib 52 are installed, localized thickness reinforcements are designed on the actual outer surface Ls relative to the theoretical outer surface Lo to achieve localized strengthening. Correspondingly, in areas of low local stress on the skin panel 1, localized thickness reduction sections are designed on the actual outer surface Ls relative to the theoretical outer surface Lo. By designing localized thickness reinforcements in areas of high stress, the strength and stiffness of the skin panel 1 in those areas can be improved, preventing tearing or failure of the skin panel 1 during aircraft operation. By designing localized thickness reduction sections in areas of low or no stress, the weight of the composite fuselage panel can be reduced, thereby reducing the overall weight of the composite fuselage panel structure.

[0042] In the section where the composite fuselage panel connects to the wing, areas with high local stress include the area where the skin panel 1 connects to the front and rear spars of the aircraft's central wing, or the area where multiple connection structures are pre-installed on the inner surface of the skin panel 1. Areas with low local stress include the area where the skin panel 1 connects to the central portion of the central wing.

[0043] It should be understood that the formation of the aforementioned localized thickness reduction portions and localized thickness reduction portions for the skin panel 1 of the composite material fuselage panel is significantly different from the method of bonding prefabricated reinforcing members to the skin panel 1, or setting the areas of increased and decreased thickness inside the panel to achieve a localized reinforcement effect. In this application, the localized thickness reinforcement portions and localized thickness reduction portions are formed by first calculating the expected stress and required strength of a localized area of ​​the skin panel 1, then calculating the number of prepreg filaments that need to be increased and / or decreased to meet this strength requirement, and then designing internal and external layering. Furthermore, after filament laying and molding using the aforementioned design, the final product is a skin panel 1 with localized thickness reinforcement portions and / or localized thickness reduction portions, i.e., the localized thickness reinforcement portions and / or localized thickness reduction portions are integrally formed with the skin panel 1.

[0044] By designing the aforementioned internal and external delaminations on the skin panel 1—that is, designing the inner surface No in a local area to have the same thickness as the theoretical outer surface Lo, and designing a local thickness reinforcement or local thickness reduction section on the side of the theoretical outer surface Lo in the corresponding local area to form the actual outer surface Ls—the inner surface quality of the composite fuselage panel can be improved. This reduces the complexity of the inner surface No of the composite fuselage panel in the assembly area, thereby reducing the possibility of shape measurement difficulties and gap compensation due to assembly positioning tolerances. Moreover, through the outer surface design, while ensuring that the composite fuselage panel itself meets the strength and stiffness requirements, the overall weight of the composite fuselage panel structure can be effectively reduced.

[0045] The design of the aforementioned theoretical outer surface Lo can be determined based on factors such as the expected drag on the aircraft and the shape of the aircraft components connected to the composite fuselage panels. The thickness of the inner surface No can be determined based on the predetermined connection structure and strength and stiffness requirements of this local area. The actual outer surface Ls design can be based on the calculated and evaluated stress conditions in the local area and the shape and thickness of the local thickness reinforcement or local thickness reduction section required to withstand the stress.

[0046] The composite material fuselage panel provided in this application, through the design of the segmented stringer 3, the circumferential local reinforcement of the skin panel 1 at the segmented position of the segmented stringer 3, and the thickness variation design of the inner surface No and the actual outer surface Ls of the skin panel 1, can ensure the stiffness continuity of the segmented stringer 3, the annular reinforcing frame 2 and the skin panel 1 in the intersection area. This ensures that the complex loads in the complex connection area can be effectively transferred. Furthermore, through the design of the inner and outer surfaces, while ensuring the stiffness of the skin panel 1, the assembly difficulty of the internal bonding surface on the composite material fuselage panel is reduced, the overall thickness and weight of the composite material fuselage panel are reduced, and the feasibility of the manufacturing process is improved.

[0047] For example, the composite material fuselage panel provided in this application can be used in the wing-body connection area, specifically in the area where the fuselage connects to the front and rear wing spars, the upper wing, and the center wing spar. Alternatively, the composite material fuselage panel provided in this application can also be used in areas of the fuselage structure covered by fairings where shape requirements are not strict.

[0048] While specific embodiments of the present invention have been described above, those skilled in the art should understand that these are merely illustrative examples, and the scope of protection of the present invention is defined by the appended claims. Those skilled in the art can make various changes or modifications to these embodiments without departing from the principles and essence of the present invention, but all such changes and modifications fall within the scope of protection of the present invention.

Claims

1. A composite material fuselage panel for an aircraft, the composite material fuselage panel comprising a skin panel, a plurality of annular reinforcing frames distributed along the heading direction of the aircraft on the inner surface of the skin panel, and a plurality of segmented stringers distributed along the circumferential direction of the aircraft between adjacent annular reinforcing frames, wherein, The inner surface of the skin panel has a circumferential local reinforcement of predetermined thickness at the intersection of the segmented truss and the annular reinforcing frame. The end of the circumferential local reinforcement extending along the heading direction overlaps with the end of the segmented truss near the circumferential local reinforcement.

2. The composite material fuselage panel according to claim 1, wherein, In the region where the circumferential local reinforcement overlaps with the end of the segmented stringer, the upper surface of the circumferential local reinforcement and the lower surface of the segmented stringer are in contact with each other.

3. The composite material fuselage panel according to claim 1, wherein, The shape of the composite fuselage panel in the area where it connects with the wing of the aircraft is designed to substantially match the shape of the area where the wing contacts it.

4. The composite material fuselage panel according to claim 1, wherein, The inner surfaces of multiple connecting areas pre-set inside the skin panel are designed to have the same thickness as the theoretical outer surface of the composite material fuselage panel, and local thickness reinforcement or local thickness reduction parts are designed on one side of the theoretical outer surface according to the local stress magnitude of the local area.

5. The composite material fuselage panel according to claim 4, wherein, The inner surface is designed as a flat inner surface at a predetermined distance from the theoretical outer surface of the composite material fuselage panel.

6. The composite material fuselage panel according to claim 5, wherein, The flat inner surface of the skin panel is designed to be either flat or curved.

7. The composite material fuselage panel according to claim 4, wherein, The local thickness reinforcement is designed on one side of the theoretical outer surface of the area with high local stress on the skin panel.

8. The composite material fuselage panel according to claim 7, wherein, The areas with high local stress include the areas where the skin panel is connected to the front and rear spars of the central wing of the aircraft, or the areas on the inner surface of the skin panel where multiple connection structures are predetermined.

9. The composite material fuselage panel according to claim 4, wherein, The localized thickness reduction section is designed on one side of the theoretical outer surface of the area where the local stress is small on the skin panel.

10. The composite material fuselage panel according to claim 9, wherein, The area with low local stress includes the region where the skin panel connects to the middle portion of the central wing.

11. The composite material fuselage panel according to any one of the preceding claims, wherein, The composite fuselage panels are configured for the wing-body connection area.

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