Composite material fuselage wall plate for aircraft

Abstract

The invention provides a composite fuselage wall plate for an aircraft, which comprises a skin wall plate, a plurality of annular reinforcing frames distributed on the inner surface of the skin wall plate along the heading direction of the aircraft, and a plurality of segmented stringers positioned between the adjacent annular reinforcing frames and distributed along the annular direction of the aircraft, an annular local reinforcing part with a preset thickness is formed at the intersection position of the segmented stringer and the annular reinforcing frame on the inner surface of the skin wallboard, and the end part, extending in the heading direction, of the annular local reinforcing part is overlapped with the end part, close to the annular local reinforcing part, of the segmented stringer.

Description

Technical Field

[0001] The present invention relates to the field of aviation, mainly to the field of composite material structure design of aircraft, and in particular to a composite material fuselage panel for aircraft. Background Art

[0002] In aircraft fuselage panels primarily made of composite materials, due to the diverse connection types and complex stress distribution in some areas of the panels, existing technologies often employ large-scale circumferential reinforcement frames and longitudinal reinforcement beams on the inner surfaces of composite fuselage panels. However, this reinforcement method adversely affects the continuity of longitudinal components, such as conventional stringers, on the fuselage skin panels, preventing the 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 a large number of annular reinforcement frames or heading reinforcement beam structures are arranged on the existing composite fuselage panels, which makes it inconvenient to install ordinary long spars on their skin panels, and proposes a new composite fuselage panel for aircraft.

[0004] The present invention solves the above technical problems through the following technical solutions:

[0005] Specifically, the present invention provides a composite fuselage panel for an aircraft, the composite fuselage panel comprising a skin panel, a plurality of annular reinforcement frames distributed on the inner surface of the skin panel along the heading direction of the aircraft, and a plurality of segmented long stringers located between adjacent annular reinforcement frames and distributed along the circumferential direction of the aircraft, wherein an annular local reinforcement portion with a predetermined thickness is formed on the inner surface of the skin panel at a position where the segmented long stringers and the annular reinforcement frames intersect, and an end portion of the annular local reinforcement portion extending along the heading direction overlaps with an end portion of the segmented long stringer close to the annular local reinforcement portion.

[0006] The composite fuselage panel for an aircraft provided in the present application combines the structural design of an annular reinforcement frame and a segmented long stringer, and connects the segmented long stringer to the skin panel by fasteners. This not only avoids the interference of the annular reinforcement frame with the installation of the segmented long stringer on the skin panel, thereby realizing the effective connection of the segmented long stringer and the annular reinforcement frame on the skin panel of the composite fuselage panel, but also comprehensively considers the restrictions of various requirements such as assembly, load transfer, stiffness and sealing at the intersection of the segmented long stringer and the annular reinforcement frame.

[0007] By designing an annular local reinforcement at the intersection of the segmented long girder and the annular reinforcement frame of the skin wall panel, and making the annular local reinforcement overlap with the end of the segmented long girder, the stiffness of the annular reinforcement, segmented long girder and skin wall panel in the intersection area can be ensured to be continuous, and the interruption of the azimuth load transmitted by the segmented long girder can be avoided, thereby achieving continuous transmission of the azimuth load on the composite fuselage wall panel and significantly improving the sealing of the composite fuselage wall panel.

[0008] According to one embodiment of the present invention, in the overlapping region between the ends of the circumferential localized reinforcement and the segmented stringers, the upper surface of the circumferential localized reinforcement and the lower surface of the segmented stringers are in contact with each other. By designing the upper and lower contact surfaces of the circumferential localized reinforcement and the segmented stringers to be in contact with each other, the stiffness of the composite fuselage panel can be maintained, the stress acting on the composite fuselage panel can be evenly distributed across the segmented stringers and the circumferential localized reinforcement, and the continuity of the load transfer in the composite fuselage panel, both in the heading direction and in the direction perpendicular to the heading load, can be achieved.

[0009] According to one embodiment of the present invention, the composite fuselage panel's profile at the area where it connects to the aircraft's wing is designed to substantially match the profile of the wing's contact area. Designing the composite fuselage panel's profile to match the profile of the wing's contact area, i.e., the profile of the wing's upper surface, facilitates connection of the composite fuselage panel to the aircraft's wing, thereby facilitating assembly of the fuselage and wing structures.

[0010] According to one embodiment of the present invention, the inner profiles of multiple connection areas pre-installed within the skin panel are designed to have the same thickness as the theoretical outer profile of the composite fuselage panel, and local thickness reinforcements or local thickness reductions are designed on the side of the theoretical outer profile based on the local stress levels in the local areas. By combining internal and external drop layers to connect and assemble a large number of areas on the composite fuselage panel, the complexity of the inner profile of the composite fuselage panel in the assembly area is reduced, thereby reducing the possibility of difficult-to-measure and difficult-to-compensate gaps due to assembly positioning tolerances between the reinforcement frame, beams, and panel.

[0011] According to one embodiment of the present invention, the inner profile is designed as a flat inner surface that is a predetermined distance from the theoretical outer profile of the composite fuselage panel. By designing the inner profile of the composite fuselage panel, which has a large number of connection structures and complex connection relationships, as a flat inner surface, significant variations in the thickness of the composite fuselage panel's inner surface or inner profile can be avoided. This allows the connection structures to be securely attached to the composite fuselage panel, thereby ensuring effective connection between the composite fuselage panel and other structural components. Furthermore, by ensuring the inner profile is at a predetermined distance from the theoretical outer profile, the basic structural strength and rigidity of the skin panel can be maintained, preventing the skin panel from failing during use.

[0012] According to one embodiment of the present invention, the flat inner surface of the skin panel is designed to be flat or curved. By designing the flat inner surface as a flat or curved surface, it can be matched with the surface with which the connecting structure contacts it. For example, if the surface with which the connecting structure contacts it is a curved surface, the flat inner surface has a curvature consistent with that of the curved surface. This facilitates movement or rapid positioning of the connecting structure on the flat inner surface of the skin panel and reduces the difficulty of assembling the connecting structure on the mating surface of the skin panel.

[0013] According to one embodiment of the present invention, a localized thickness reinforcement is designed on one side of a theoretical profile in an area of ​​high local stress on a skin panel. By designing the required localized thickness reinforcement on one side of the theoretical profile, the resulting localized thickness reinforcement constitutes a portion of the formed skin panel (or, in other words, the reinforcement or the reinforcement area itself is located within the skin panel). Therefore, compared to pre-defined reinforcement sheets bonded to the outer surface of the skin panel, this has advantages such as more reliable manufacturing quality, lower cost, and better mechanical properties of the joint surface.

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

[0015] According to one embodiment of the present invention, a localized thinning section is designed on one side of the theoretical profile surface of the skin panel in an area of ​​low local stress. By designing the thinning section in a localized area of ​​low or minimal stress, the weight of the composite 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 a region where the skin panel is connected to the middle portion of the center wing.

[0017] According to one embodiment of the invention, a composite fuselage panel is configured for the wing-to-fuselage connection area.

[0018] On the basis of conforming to the common sense in this field, the above preferred embodiments can be arbitrarily combined to obtain the preferred embodiments of the present invention.

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

[0020] 1. For composite fuselage panels with annular reinforcement frames, segmented long stringers are used to absorb and transfer azimuth loads, avoiding interference of the annular reinforcement frames with the assembly of the long stringers, and enabling the segmented long stringers to be firmly fixed to the skin panel through mechanical connection. At the same time, at the long stringer segments where the azimuth load is greater, a circumferential local reinforcement with a certain thickness and width covering the end of the segmented long stringers is designed on the inner surface of the skin panel. This can achieve effective connection between the segmented long stringers and between the segmented long stringers and the skin panel and the annular reinforcement frame, ensuring the continuity of stiffness of the annular reinforcement frame, skin panel and segmented long stringers in the intersection area, thereby achieving the connection and transfer of azimuth loads and the complex load transfer between various components or structures;

[0021] 2. By adopting a uniform internal profile design across numerous joint assembly areas on the skin panels, the complexity of the composite panel internal profile in these areas can be reduced, thereby minimizing the likelihood of difficult-to-measure and gap-compensating shapes due to axial or circumferential assembly positioning tolerances. Furthermore, the assembly difficulty of the butt-joint structural surfaces is reduced, reducing the number of required compensation shims. This effectively improves the structural integrity of the composite fuselage panels, reduces assembly process complexity, and increases overall assembly speed.

[0022] 3. Moreover, by designing the internal and external layer loss on the skin panel, that is, designing the inner surface of the local area to be the same thickness, and designing the local thickness reinforcement part or the local thickness thinning part on the side of the outer surface of the corresponding local area, it is possible to ensure that the composite fuselage panel itself meets the strength and stiffness requirements while effectively reducing the overall weight of the composite fuselage panel structure. BRIEF DESCRIPTION OF THE DRAWINGS

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

[0024] Figure 2 Schematically shows Figure 1 Cutaway cross-sectional view of the composite fuselage panel along direction A.

[0025] Figure 3 A schematic diagram of a portion of a composite fuselage panel arranged with a large number of connections is shown.

[0026] Figure 4 Schematically shows Figure 3Cross-sectional view of the composite fuselage panel along direction B. DETAILED DESCRIPTION

[0027] To make the objectives, technical solutions, and advantages of this application more clear, the following will clearly and completely describe the technical solutions in the embodiments of this application in conjunction with the accompanying drawings showing multiple embodiments according to this application. It should be understood that all other embodiments obtained by ordinary technicians in this field based on the embodiments described in this application without expending 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 meanings as commonly understood by those skilled in the art to which this application belongs. The terms "including" and "having" in the specification and claims of this application and the above-mentioned description of the drawings are open-ended terms. Therefore, "including" and "having" are, for example, one or more parts or portions, and it has one or more parts or portions, but are not limited to having only these one or more parts or portions.

[0029] In the aviation field, composite materials are widely used in aircraft fuselage structures, particularly as the primary material for fuselage panels. Aircraft fuselage panels often require numerous complex joint structures, necessitating complex interface structures and a certain degree of rigidity and strength to meet these requirements.

[0030] In order to ensure the strength of the composite fuselage panels, annular metal reinforcement frames distributed along the aircraft's heading direction and aircraft long stringers or heading reinforcement beams distributed along the aircraft's circumferential direction are often arranged on the skin panels of the composite fuselage panels to ensure the strength of the formed composite fuselage panels so that they can withstand loads from other aircraft components or impact forces from the flight environment.

[0031] However, the large number of large, circular metal reinforcement frames arranged on the composite fuselage panels, particularly at the fuselage-wing junction, the most complex area of ​​the aircraft fuselage structure, can affect the installation of navigation components, including stringers, on the panels. This can prevent the stringers from being securely mounted on the panels or create gaps between the stringers and the panels, making the panels susceptible to tearing or failure during use. Furthermore, according to engineering experience, this cross-connection of the circular metal reinforcement frames with the stringers often results in stress concentration at the intersection of the stringers and reinforcement frames, and can also lead to defects such as discontinuous load transfer between the composite fuselage panels.

[0032] Therefore, the present application provides a new composite fuselage panel for aircraft. Figure 1 As shown, the composite fuselage panel comprises a skin panel 1, a plurality of annular reinforcement frames 2 distributed along the aircraft's heading direction on the inner surface of the skin panel 1, and a plurality of segmented long stringers 3 distributed along the aircraft's circumferential direction between adjacent annular reinforcement frames 2. Figure 2 As shown, the inner surface of the skin panel 1 is provided with an annular local reinforcement portion 11 with a predetermined thickness at the segmented position of the segmented long stringer 3 where the heading load is larger, that is, the position where the segmented long stringer 3 intersects the annular reinforcement frame 2, and the end of the annular local reinforcement portion 11 extending along the heading direction overlaps with the end of the segmented long stringer 3 close to the annular local reinforcement portion 11.

[0033] like Figure 2 As shown, in the overlapping area between the end of the circumferential local reinforcement 11 and the segmented long stringer 3, the upper surface of the circumferential local reinforcement 11 and the lower surface of the segmented long stringer 3 are in contact with each other to ensure the rigidity of the composite fuselage panel. The stress on the composite fuselage panel can be evenly distributed on the segmented long stringer 3 and the circumferential local reinforcement 11. The heading load is then transferred to the adjacent segmented long stringer 3 via the circumferential local reinforcement 11, and the load perpendicular to the heading load direction is transferred to the annular reinforcement frame 2 and the skin panel 1. In this way, the continuity of the heading load transfer and the load transfer perpendicular to the heading load direction on the composite fuselage panel is achieved.

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

[0035] In terms of appearance design, when composite fuselage panels are used to connect with wing structures, such as Figure 3 As shown, the composite fuselage panel's profile at the area where it connects to the aircraft's wing is designed to substantially match the profile of the wing's contact area. This profile design enables form fit and effective connection between the composite fuselage panel and the aircraft's wing, thereby enabling assembly of the fuselage and wing structures.

[0036] The area where the fuselage and wing structures meet is one of the most complex areas in aircraft structure. Specifically, the composite fuselage skin panels (1) exhibit complex connections and significant stress variations. Existing techniques typically employ methods such as bonding prefabricated reinforcements to corresponding sides of the skin panels (1), or implementing thickness variations at the connection area of ​​the panel's theoretical inner profile to achieve localized reinforcement.

[0037] When a thickness variation design is performed at a large number of connection areas of the theoretical inner profile for local reinforcement, the inner profile at the skin panel connection area will have thickness variations along the circumferential direction. The connection structure (such as the circumferential reinforcement) that fits the skin panel in this area will produce a circumferential displacement relative to the theoretical position during installation. Therefore, a large number of gaps will be generated between the inner profile of the skin panel and the connection structure in this area. In this case, in order to ensure the fit and connection of the connection structure on the skin panel, a large number of gaskets with a parallelogram cross-section at two locations are required to compensate for the gaps, which increases cost and cannot ensure a stable connection between the connection structure and the skin panel. More seriously, when a thickness variation design is performed at the connection area of ​​the theoretical inner profile for local reinforcement, the inner profile at the skin panel connection area will have thickness variations along the circumferential direction and the axial direction. At this time, when the circumferential reinforcement or the connection structure fits the skin panel in this area, it will be offset in the circumferential direction and the axial direction relative to the theoretical position. In this case, complex gaps may exist between the connecting structure mounted on the skin panel and the panel, necessitating the use of complex shims with cross-sections in different directions, such as a combination of wedges and trapezoids, to compensate for the gaps. This not only requires considerable time and expense to manufacture the complex shims required to compensate for the gaps, but also makes it difficult to effectively fill the gaps and achieve an accurate and effective connection between the skin panel and the connecting structure.

[0038] As previously mentioned, when the circumferential and longitudinal reinforcement structures exhibit misalignment, complex gaps can easily form between the panels, making them difficult to measure and effectively fill. However, due to the complex internal force or stress distribution in areas with complex connections, localized reinforcement using prefabricated reinforcements bonded to one side is difficult to implement and its quality is also difficult to ensure. Extensive thickness variations within the theoretical inner profile of the panel can make it difficult for the connection structure to conform to the surface of the skin panel 1 for assembly onto the composite fuselage panel.

[0039] In response to the above problems, Figures 3 to 4 As shown, the composite fuselage panel provided in the present application also uses its theoretical outer profile Lo as a reference surface, designs the inner profile No at a large number of predetermined connection areas within the skin panel 1 to have a uniform thickness, and designs local thickness reinforcements or local thickness reductions on the side of the theoretical outer profile Lo based on the magnitude of local stress in the local areas to form an actual outer profile Ls. Therefore, the inner profile No at the multiple connection areas predetermined within the skin panel 1 is designed to have a uniform thickness relative to the theoretical outer profile Lo of the composite fuselage panel, and the actual outer profile Ls at the local areas where the local stress is high or low is designed to have local thickness reinforcements or local thickness reductions relative to its theoretical outer profile Lo.

[0040] Specifically, using the theoretical outer profile Lo of the composite fuselage panel as a reference surface, the inner profile No is designed to be a flat inner surface at a predetermined distance from the reference surface. The shape of this flat inner surface is designed to be either plane or curved to match the shape of the surface with which the connecting structure is intended to be mounted. For example, if the surface with which the connecting structure contacts is a curved surface, the flat inner surface may have a curvature consistent with that of the curved surface. This facilitates the movement or rapid positioning of the connecting structure on the flat inner surface of the skin panel and reduces the difficulty of assembling the connecting structure on the mating surface of the skin panel.

[0041] In addition, a local thickness reinforcement portion is designed on one side of the theoretical outer surface Lo at the area with large local stress on the skin panel 1, that is, the portion between the actual outer surface Ls and the theoretical outer surface Lo. Figure 4 As shown, a local thickness reinforcement portion is designed on the actual outer surface Ls of the local area where the complex connection structure such as the central wing joint 51 and the central wing No. 1 rib edge strip 52 is installed on the skin panel 1, relative to the theoretical outer surface Lo, for achieving a local reinforcement effect. Correspondingly, a local thickness thinning portion is designed on the actual outer surface Ls of the skin panel 1 in the area where the local stress is small, relative to the theoretical outer surface Lo. By designing a local thickness reinforcement portion in a local area where the stress is large, the strength and stiffness of the skin panel 1 in this local area can be improved, thereby preventing the skin panel 1 from tearing or failing during the operation of the aircraft. By designing a local thickness thinning portion in a local area where the stress is low or the stress is small, the weight of the composite fuselage panel can be reduced, thereby reducing the overall weight of the composite fuselage panel structure.

[0042] In the area where the composite fuselage panels connect to the wings, areas with high local stress include where the skin panel 1 connects to the front and rear beams of the aircraft's center wing, or where multiple connection structures are pre-determined on the inner surface of the skin panel 1. Areas with low local stress include where the skin panel 1 connects to the middle portion of the center wing.

[0043] It should be understood that for the skin panel 1 of the composite fuselage panel, the formation of the above-mentioned local thickness reduction portion and the local thickness reduction portion is significantly different from the method of bonding prefabricated reinforcements to the skin panel 1, or arranging the areas of increased and reduced thickness inside the panel to achieve the local reinforcement effect. The local thickness reinforcement portion and the local thickness reduction portion in the present application are first calculated by first calculating the stress and required strength that the local area of ​​the skin panel 1 is expected to withstand, and then calculating the number of prepreg bundles that need to be increased and / or reduced to meet the strength requirements, and then performing the design of internal and external layer loss. Moreover, after the aforementioned design is used for wire laying and molding, what is finally formed is a skin panel 1 with a local thickness reinforcement portion and / or a local thickness reduction portion, that is, the local thickness reinforcement portion and / or the local thickness reduction portion are integrally molded with the skin panel 1.

[0044] By designing the aforementioned internal and external layer dropouts on the skin panel 1, specifically designing the inner profile No in a local area to have the same thickness as the theoretical outer profile Lo, and designing a local thickness reinforcement or thinning portion on the side of the theoretical outer profile Lo in the corresponding local area to form the actual outer profile Ls, the quality of the inner surface of the composite fuselage panel can be improved, thereby reducing the complexity of the inner profile No of the composite fuselage panel in the assembly area, and thus reducing the possibility of difficult shape measurement and gap compensation due to assembly positioning tolerances. Furthermore, the outer profile design can effectively reduce the overall weight of the composite fuselage panel structure while ensuring that the composite fuselage panel itself meets the strength and stiffness requirements.

[0045] The design of the theoretical outer profile Lo can be determined based on factors such as the expected drag on the aircraft and the external shapes of the aircraft components to which the composite fuselage panel is attached. The thickness of the inner profile No can be determined based on the predetermined connection structure, strength, and stiffness requirements for the local area. The actual outer profile Ls can be designed based on the calculated and assessed stress conditions in the local area and the shape and thickness of the local thickness reinforcement or local thickness reduction required to withstand these stresses.

[0046] The composite fuselage panel provided in the present application can ensure the continuity of stiffness of the segmented long stringers 3, the annular reinforcement frame 2 and the panel skin 1 in the intersection area through the design of the segmented long stringers 3, the circumferential local reinforcement portion of the skin panel 1 at the segmented position of the segmented long stringers 3, and the thickness variation design of the inner profile No and the actual outer profile Ls of the skin panel 1, thereby ensuring that complex loads in the complex connection relationship area can be effectively transmitted. It can also reduce the assembly difficulty of the internal fitting surface of the composite fuselage panel while ensuring the stiffness of the skin panel 1 through the design of the inner and outer profiles, reduce the overall thickness and weight of the composite fuselage panel, and improve the feasibility of the manufacturing process.

[0047] For example, the composite fuselage panels provided herein can be used in the wing-to-body connection area, specifically in the areas where the fuselage connects to the front and rear wing spars, the wing top, and the center wing strip. Alternatively, the composite fuselage panels provided herein can also be used in areas of the fuselage structure covered by fairings and where the appearance requirements are less stringent.

[0048] Although specific embodiments of the present invention have been described above, those skilled in the art will appreciate that these are merely illustrative and that the scope of the present invention is defined by the appended claims. Those skilled in the art may make various changes or modifications to these embodiments without departing from the principles and essence of the present invention, and such changes and modifications are intended to fall within the scope of the present invention.

Claims

1. A composite fuselage panel for an aircraft, comprising a skin panel, a plurality of annular reinforcement frames distributed on the inner surface of the skin panel along the heading direction of the aircraft, and a plurality of segmented stringers located between adjacent annular reinforcement frames and distributed along the circumferential direction of the aircraft, wherein: An annular local reinforcement portion with a predetermined thickness is formed on the inner surface of the skin wall panel at the position where the segmented long stringer intersects the annular reinforcement frame, and the end of the annular local reinforcement portion extending along the heading direction overlaps with the end of the segmented long stringer close to the annular local reinforcement portion.

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

3. The composite fuselage panel according to claim 1, wherein: The profile of the composite fuselage panel at the region where it connects to the wing of the aircraft is designed to substantially match the profile of the wing at the region where it contacts the wing.

4. The composite fuselage panel according to claim 1, wherein: The inner profiles of multiple connection areas pre-arranged inside the skin panel are designed to have the same thickness relative to the theoretical outer profile of the composite fuselage panel, and local thickness reinforcement portions or local thickness thinning portions are designed on one side of the theoretical outer profile according to the magnitude of local stress in the local area.

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

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

7. The composite fuselage panel according to claim 4, wherein: The local thickness reinforcement portion is designed on one side of the theoretical outer surface of the region where the local stress of the skin panel is large.

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

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

10. The composite fuselage panel according to claim 9, wherein: The region with low local stress includes the region where the skin panel is connected to the middle portion of the center wing.

11. A composite fuselage panel according to any one of the preceding claims, wherein The composite fuselage panel is configured for use in the wing-to-fuselage connection area.

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