Skin assembly of aircraft cylindrical section, and frame of aircraft cylindrical section

Abstract

A skin assembly of an aircraft cylindrical section (100). The skin assembly comprises a plurality of wallboards which are sequentially connected in a circumferential direction and have flush inner surfaces, wherein the inner surfaces of all the wallboards form a smooth cylindrical surface; the outer surface of each wallboard comprises a plurality of equal-thickness areas (101) and transition areas (102) which are alternately arranged in a heading direction; the thickness of any equal-thickness area (101) is not less than a corresponding minimum load thickness; the surface of any transition area (102) is an annular surface arranged in the circumferential direction; the boundaries of two sides of any transition area (102) are respectively in contact with adjacent equal-thickness areas (101), with the thicknesses of the boundaries of the two sides of the transition area (102) being respectively equal to the thicknesses of the equal-thickness areas (101) in contact therewith; and any transition area (102) has a preset slope in the heading direction. A frame of the aircraft cylindrical section (100), the frame comprising the skin assembly of the aircraft cylindrical section (100). Such arrangements ensure the aerodynamic performance of an aircraft and also simplify the configurations of stringers and circumferential frames fitting closely onto the inner surface of an aircraft skin, thereby being conducive to reducing the manufacturing costs of the aircraft.

Description

Straight skin components and straight frames of aircraft, etc. Technical Field

[0001] This application belongs to the field of aviation manufacturing technology and relates to an aircraft skin structure, particularly to a straight skin assembly and a straight frame of an aircraft. Background Technology

[0002] Composite materials are materials composed of multiple materials with different properties or different organizational forms, and are a commonly used material for aircraft skin. In order to improve the structural efficiency of aircraft skin, that is, to reduce weight as much as possible while ensuring the strength and rigidity of the aircraft skin, thereby reducing costs and improving performance, it is necessary to remove layers from the laminated structure of the aircraft skin. That is, the number of layers in different locations of the aircraft skin is different, resulting in uneven thickness between different areas of the aircraft skin.

[0003] In existing technologies, to ensure aircraft aerodynamic performance, the aircraft skin is layered inwards, meaning that some composite material layers are selectively reduced on the inner side of the skin. This results in a smooth outer surface, uniform airflow distribution, and avoids vortex formation. However, this structural design also makes the inner surface of the skin uneven, leading to a more complex configuration of the stringers and circumferential frames that are tightly attached to the inner side of the skin. This increases the difficulty of manufacturing and assembly, resulting in higher aircraft manufacturing costs.

[0004] Therefore, how to design an aircraft skin that simplifies the configuration of the stringers and circumferential frames while ensuring the aerodynamic performance of the aircraft is a problem that urgently needs to be solved by those skilled in the art. Summary of the Invention

[0005] The purpose of this application is to provide a straight-section skin assembly and a straight-section frame for aircraft, etc., to solve the problem that existing aircraft skins, due to their inward-layering structural design, have uneven inner surfaces, resulting in complex configurations of the stringers and circumferential frames, high manufacturing and assembly difficulties, and consequently, high manufacturing costs.

[0006] In a first aspect, this application provides a straight-section skin assembly for an aircraft, including multiple circumferentially connected panels with flush inner surfaces. The inner surfaces of all the panels form a smooth cylindrical surface. The outer surfaces of each panel include multiple alternating equal-thickness regions and transition regions in the flight direction. The thickness of any equal-thickness region is not less than the corresponding minimum load thickness. The surface of any transition region is a toroidal surface arranged in the circumferential direction. The two side boundaries of any transition region are in contact with the adjacent equal-thickness regions, and the thickness at the two side boundaries of the transition region is equal to the thickness of the equal-thickness regions it contacts. Each transition region has a preset slope in the flight direction.

[0007] The minimum load thickness is the minimum thickness of the equal-thickness region that can support the maximum load at the corresponding position.

[0008] In this application, a smooth cylindrical surface is formed by the inner surface of all the wall panels. Without affecting the aerodynamic performance of the aircraft, the stringers and circumferential frames that are close to the inner surface of the straight skin components of the aircraft are simple in configuration, with fewer parts and simpler assembly. This reduces the manufacturing cost and assembly difficulty of the aircraft, which is beneficial to the actual production and manufacturing of the aircraft.

[0009] In one embodiment of this application, the preset slope is the slope at which the wind resistance of the straight part of the aircraft is within a preset wind resistance range.

[0010] In one embodiment of this application, the preset slope is 1:100.

[0011] In one embodiment of this application, the thickness difference between the equal-thickness regions with overlapping portions at any two boundaries is no greater than a preset difference threshold.

[0012] In one embodiment of this application, the preset difference threshold is 1 mm.

[0013] In one embodiment of this application, if the difference in minimum load thickness between any two equal-thickness regions with overlapping boundaries is not greater than the preset difference threshold, then the thickness of the equal-thickness region is the corresponding minimum load thickness.

[0014] In one embodiment of this application, four wall panels are spliced ​​together and have flush inner surfaces, wherein the wall panel corresponding to the left side of the straight part of the aircraft is mirror-symmetrical to the wall panel corresponding to the right side of the straight part of the aircraft.

[0015] Secondly, this application also provides a straight frame for an aircraft, including the straight skin assembly of the aircraft as described above, as well as a plurality of stringers and a plurality of circumferential frames. The stringers are structures extending along the flight direction of the straight part of the aircraft, and the circumferential frames are annular structures that are closed along the circumferential direction of the straight part of the aircraft. All the stringers and all the circumferential frames are disposed in contact with the inner surface of the straight skin assembly of the aircraft.

[0016] In one embodiment of this application, a fairing is further included for improving the smoothness of the straight outer surface of an aircraft or similar device, the fairing at least covering the connection between the straight skin assembly of the aircraft or similar device and the external structure.

[0017] In one embodiment of this application, a sealant is also included. The straight skin assembly of the aircraft or the like comprises multiple interlocking wall panels. The sealant covers the contact boundary on the outer surfaces of adjacent wall panels to improve the smoothness of the outer surface of the straight part of the aircraft or the like.

[0018] As described above, this application provides a straight-section skin assembly and a straight-section frame for an aircraft. By using multiple interconnected wall panels with flush inner surfaces, the inner surface of the straight-section skin assembly is made into a smooth cylindrical surface, thereby simplifying the configuration of the stringers and circumferential frames, effectively reducing the manufacturing cost and assembly difficulty of the stringers and circumferential frames, and thus reducing the production cost of the aircraft. Furthermore, by using multiple equal-thickness regions and transition regions alternately arranged along the flight direction on the outer surface of each wall panel, the load requirements of the aircraft in different regions are met while ensuring that the thickness of each wall panel is constant in the circumferential direction, thereby giving the aircraft good aerodynamic performance. This ensures that the inner surface of the straight-section skin assembly is smooth, reducing the production cost of the aircraft while improving its flight performance, which is beneficial for the actual production and application of the aircraft. Attached Figure Description

[0019] Figure 1 shows a schematic diagram of a composite material delaminated structure according to an embodiment of this application.

[0020] Figure 2 shows a schematic diagram of the main structure of an aircraft according to an embodiment of this application.

[0021] Figure 3 shows a schematic diagram of the straight structure of an aircraft according to an embodiment of this application.

[0022] Figure 4 shows a schematic diagram of an explosion of a straight structure such as an aircraft according to an embodiment of this application.

[0023] Figure 5 shows a schematic diagram of an aircraft wall panel structure according to an embodiment of this application.

[0024] Figure 6 shows a schematic diagram of another aircraft panel structure according to an embodiment of this application.

[0025] Figure 7 shows a schematic diagram of the airflow velocity distribution of an aircraft body according to an embodiment of this application.

[0026] Component designation: 100 - Straight section of the aircraft; 110 - Upper panel; 120 - Right panel; 130 - Left panel; 140 - Lower panel; 101 - Region of uniform thickness; 102 - Transition region; 200 - Nose of the aircraft; 300 - Tail of the aircraft. Detailed Implementation

[0027] The following specific examples illustrate the implementation of this application. Those skilled in the art can easily understand other advantages and effects of this application from the content disclosed in this specification. This application can also be implemented or applied through other different specific embodiments, and various details in this specification can also be modified or changed based on different viewpoints and applications without departing from the spirit of this application. It should be noted that, unless otherwise specified, the following embodiments and features in the embodiments can be combined with each other.

[0028] It should be noted that the illustrations provided in the following embodiments are only schematic representations of the basic concept of this application. Therefore, the drawings only show the components related to this application and are not drawn according to the actual number, shape and size of the components in the actual implementation. In the actual implementation, the form, quantity and proportion of each component can be arbitrarily changed, and the layout of the components may also be more complex.

[0029] To improve the structural efficiency of aircraft skin, it is necessary to reduce the number of layers in the composite material laminate structure. Specifically, as shown in Figure 1, the required thickness of the aircraft skin varies depending on the load at different locations on the aircraft. By reducing the number of composite material layers in areas with lower loads, the weight of the aircraft can be effectively reduced, thereby lowering costs and improving performance. However, this design results in unevenness on one side of the composite material surface, leading to an uneven aircraft skin. In existing technologies, to prevent the unevenness of the aircraft shell from affecting the gas velocity distribution on the aircraft surface and thus impacting aerodynamic performance, this uneven surface is generally used as the inner surface of the aircraft skin, resulting in a smooth curved outer surface. However, this aircraft skin structure requires matching with complex stringers and circumferential frames to ensure they fit snugly against the skin, making the manufacturing of the stringers and circumferential frames difficult and assembly challenging, thus leading to higher manufacturing costs.

[0030] Among them, the aircraft includes, but is not limited to, aerospace vehicles such as airplanes or rockets. The inner surface of the aircraft skin is the surface of the aircraft skin facing the internal space of the aircraft, and the outer surface of the aircraft skin is the surface of the aircraft skin facing the external space of the aircraft.

[0031] The following embodiments of this application provide a straight-section skin assembly and a straight-section frame for an aircraft. By using the flat side of the composite material as the inner surface of the straight-section skin assembly to form a smooth cylindrical surface, and the transition area of ​​the outer surface has a preset slope, the aerodynamic performance of the aircraft is guaranteed while eliminating the need for complex configurations of the stringers and circumferential frames to fit the aircraft skin. This simplifies the manufacturing and assembly of the stringers and circumferential frames, effectively reduces the manufacturing cost of the aircraft, and solves the problem of high manufacturing and assembly difficulty of the stringers and circumferential frames in the prior art.

[0032] The following will describe in detail, with reference to the accompanying drawings, the principle and implementation method of a straight-section skin assembly and a straight-section frame of an aircraft, etc., according to this embodiment, so that those skilled in the art can understand the straight-section skin assembly and the straight-section frame of an aircraft, etc., according to this embodiment without creative effort.

[0033] As shown in Figure 1, this embodiment provides a skin assembly. In order to improve the structural efficiency of the aircraft, the aircraft skin assembly has different thicknesses for different load areas of the aircraft, that is, one side of the aircraft skin assembly is an uneven surface.

[0034] Figure 2 shows a schematic diagram of the main body of the aircraft. The main body includes a nose section 200, a vertical section 100, and a tail section 300. The nose section 200 is located at the front of the aircraft, and the tail section 300 is located at the rear. The nose section 200 and the tail section 300 are connected by the vertical section 100. The nose section 200, the vertical section 100, and the tail section 300 are all constructed from composite materials to form the aircraft shell. In this application, the uneven side of the composite material is used as the outer surface of the aircraft skin assembly, thereby simplifying the configuration of the stringers and circumferential frames that fit the inner surface of the aircraft skin assembly. This reduces the manufacturing and assembly difficulty of the stringers and circumferential frames, and consequently lowers the manufacturing cost of the aircraft.

[0035] It should be noted that, in order to reduce the drag of the aircraft, the shape of the main body of the aircraft is usually a streamlined shape with low drag. Specifically, the curvature of each point on the nose 200 of the aircraft changes smoothly, and the tail 300 is connected by a cylindrical straight section 100. For example, as shown in Figure 2, when the aircraft is an airplane, the curvature of each point on the nose 200 and the tail 300 changes smoothly, and the surface of the straight section 100 is a cylinder with constant curvature at each point. Based on this, in order to ensure the aerodynamic performance of the aircraft, the outer surface of the aircraft skin assembly needs to meet the preset flatness requirements to avoid affecting the normal flight of the aircraft. Among them, the parts with larger curvature changes are more sensitive to airflow changes, while the areas with constant curvature have less impact on aerodynamic performance. That is, the flatness requirement of the outer surface of the straight section 100 is much lower than that of the nose 200 and the tail 300.

[0036] This embodiment will take the aircraft straight section skin assembly located at position 100 of the aircraft as an example to illustrate its specific structure and working principle.

[0037] As shown in Figure 3-4, the straight-section skin assembly of an aircraft comprises multiple interconnected panels with flush inner surfaces. The inner surfaces of these panels form a smooth cylindrical surface to simplify the configuration of the stringers that fit the inner surfaces and the configuration of the circumferential frame. It should be noted that each panel can be manufactured in a one-piece molding process, and the entire straight-section skin assembly of the aircraft can be obtained by avoiding splicing the individual panels.

[0038] To reduce the weight of the aircraft, the composite laminate structure requires layer dropping to reduce weight, resulting in an uneven outer surface for these panels. Specifically, each panel's outer surface includes multiple alternating equal-thickness regions 101 and transition regions 102 along the flight direction. The equal-thickness regions 101 are areas with uniform thickness, while the transition regions 102 are all annular surfaces arranged circumferentially. Therefore, the outer surface of a single panel does not change thickness in the circumferential direction, but only in the flight direction. It should be noted that when there are thickness variations in multiple directions on the outer surface of the skin, the airflow becomes complex, significantly impacting the aircraft's aerodynamic performance and potentially creating vortices on the surface, affecting normal flight. Therefore, in this embodiment, the outer surface of each panel only has thickness variations in the flight direction, while the thickness remains constant in the circumferential direction. This avoids thickness variations in multiple directions on the aircraft's outer surface, thereby minimizing the impact of surface unevenness on the aircraft's aerodynamic performance. Based on this, in this embodiment, the thickness of each panel does not change in the circumferential direction, that is, the thickness of each panel is constant in the circumferential direction, and the thickness is only changed in the heading direction. This allows each panel to change its thickness based on the load, so as to ensure the structural efficiency of each region while avoiding the impact on the aerodynamic performance of the aircraft, thereby obtaining better flight performance. In addition, since the inner surface of the aircraft skin is smooth and the stringer and circumferential frame configuration is simple, the manufacturing cost of the aircraft is effectively reduced, which is beneficial to the practical application of the aircraft.

[0039] Here, heading is the direction in which the aircraft flies, and circumferential direction is the circumferential direction around the central axis of the aircraft fuselage.

[0040] Furthermore, the thickness of the equal-thickness region 101 is not less than the corresponding minimum load thickness. The transition region 102 has a preset slope in the flight direction. The two side boundaries of the transition region 102 are in contact with the adjacent equal-thickness region 101, and the thickness at the two side boundaries of the transition region is equal to the thickness of the corresponding equal-thickness region 101, so as to achieve a smooth transition between the equal-thickness regions 101. The minimum load thickness is the minimum thickness of the equal-thickness region 101 that can support the maximum load at the corresponding position, that is, the thickness of each equal-thickness region 101 when it bears the maximum load at the corresponding position on the aircraft, which is the minimum load thickness of the corresponding equal-thickness region 101.

[0041] It should be noted that, as shown in Figure 1, the transition region 102 is formed due to the loss of layers in the laminated structure of the aircraft skin, and the width of the transition region 102 is much smaller than that of the equal-thickness region 101. Furthermore, as shown in Figure 5, the transition region 102 connects the boundaries of two equal-thickness regions 101. If the thickness difference between the boundaries of the two equal-thickness regions 102 is h1, that is, the thickness difference between the two sides of the transition region 102 is h1, and the width of the transition region 102 is x, then the toroidal slope of this transition region is the ratio of h1 to x, that is, the preset slope is h1:x.

[0042] It should be noted that this embodiment is used as an example for ease of understanding. The preset slope is calculated based on the thickness difference and width between the two sides of the transition region 102. However, in the actual manufacturing process of the aircraft skin, the thickness difference of the equal-thickness regions 101 on both sides of the transition region 102 is obtained as the thickness difference between the two sides of the transition region 102, and the width of the transition region 102 is obtained based on the preset slope.

[0043] Preferably, the preset slope is the slope of the vertical part 100 of the aircraft during flight, where the wind resistance is within a preset wind resistance range. The preset wind resistance range is obtained based on the flight requirements of the aircraft, and the specific method of obtaining it is common knowledge in the field, which should be known to those skilled in the art, and will not be elaborated here.

[0044] It should be noted that, since the annular surface of the transition region 102 will generate a certain amount of wind resistance during flight, affecting the flight of the aircraft, in this embodiment, the slope of the annular surface of the transition region 102 is set to meet the requirements of smooth airflow, so as to reduce the wind resistance of the aircraft during flight and further enhance the aerodynamic performance of the aircraft.

[0045] For example, the preset slope is 1:100, that is, as shown in Figure 5, h1:x is 1:100. If the thickness change of the equal thickness regions 101 on both sides is 1mm, then the width of the transition region 102 is 100mm. At this time, the thickness change between the equal thickness regions 101 is relatively smooth, and the wind resistance generated during the flight of the aircraft is very small, which will not affect the normal flight of the aircraft, thus achieving better flight performance.

[0046] In some alternative implementations, since the aircraft skin comprises multiple panels and different areas of the aircraft have load requirements, the thickness of each panel is different, and the boundaries of the contacting panels have thickness differences. In order to further simplify the configuration of the stringers and circumferential frames, in this embodiment, they are avoided from being connected to each other and their inner surfaces are flush, so that the inner surface of the straight skin assembly of the aircraft forms a smooth cylindrical surface. This simplifies the structure of the stringers and circumferential frames that fit the inner surface of the straight skin assembly of the aircraft, effectively reducing the manufacturing cost of the aircraft.

[0047] It should be noted that, as shown in Figure 6, due to the possible thickness difference at the contact boundaries of the various panels, the outer surface of the aircraft skin is not smooth. However, at any point on the contact boundaries of the various panels, the thickness difference is not greater than a preset difference threshold, so as to minimize the difference in the smoothness of the straight part of the aircraft's outer surface caused by the thickness difference between the panels, thereby further improving the aerodynamic performance of the aircraft.

[0048] It should be noted that, as shown in Figure 6, the thickness difference at the boundary where the wall panels are in contact is actually the thickness difference of the equal-thickness regions with overlapping portions at the boundaries of two adjacent wall panels, i.e., the value of h2. Based on this, to ensure that the thickness difference at any point on the boundary where the wall panels are in contact does not exceed a preset difference threshold, in some optional embodiments, for any two equal-thickness regions 101 with overlapping portions at their boundaries, the thickness difference between them does not exceed the preset difference threshold, thus achieving the goal that the thickness difference at any point on the boundary where the wall panels are in contact does not exceed the preset difference threshold.

[0049] It should be noted that, since the equal-thickness regions 101 and transition regions 102 are arranged alternately on the same wall panel, there is no overlapping boundary between any two equal-thickness regions 101 on the same wall panel. The equal-thickness regions 101 with overlapping boundaries are actually located on two adjacent wall panels. For example, as shown in Figures 3-4, the boundaries of equal-thickness region C and equal-thickness region A overlap, and similarly, the boundaries of equal-thickness region C and equal-thickness region B overlap.

[0050] For example, the preset difference threshold is 1mm, that is, for any two equal-thickness regions 101 with overlapping boundaries, the thickness difference does not exceed 1mm.

[0051] It should be noted that if the difference in minimum load thickness between any two equal-thickness regions 101 with overlapping boundaries is not greater than a preset difference threshold, then the thickness of the equal-thickness region 101 is the corresponding minimum load thickness. That is, based on the load of each region of the straight section 100 of the aircraft, the minimum load thickness of all equal-thickness regions 101 is calculated, and the difference in minimum load thickness between all equal-thickness regions 101 with overlapping boundaries is obtained. If these differences are not greater than a preset difference threshold, then the thickness of all equal-thickness regions 101 is the corresponding minimum load thickness. This reduces the thickness of all equal-thickness regions 101, thereby reducing the weight of the aircraft and improving its flight performance.

[0052] Furthermore, if the difference in minimum load thickness between any two equal-thickness regions 101 with overlapping boundaries is greater than a preset difference threshold, the maximum value of all minimum load thicknesses is obtained as the thickness of these equal-thickness regions 101 with overlapping boundaries, thereby improving the flatness of the aircraft skin and further enhancing the aerodynamic performance of the aircraft.

[0053] Alternatively, the maximum value of all minimum load thicknesses can be obtained, and this maximum value can be used as the thickness of the corresponding equal-thickness region 101. Based on a preset difference threshold, the thicknesses of other equal-thickness regions 101 can be obtained. For example, the difference between the maximum value and the preset difference threshold can be used as the thickness of equal-thickness regions 101 that overlap or partially overlap with the boundary of the equal-thickness region 101 corresponding to the maximum value, and so on, to obtain the thicknesses of all other equal-thickness regions 101. Based on this, the aerodynamic performance of the aircraft can be maintained while reducing its weight, thus enabling the aircraft to achieve good flight performance.

[0054] It should be noted that the above embodiment provides an exemplary method for obtaining the thickness of the equal-thickness region 101, but it is not limited thereto. Those skilled in the art can design equal-thickness regions 101 with different thicknesses according to actual needs, as long as the thickness of all equal-thickness regions 101 is not less than the corresponding minimum load thickness, and the difference in thickness between any two equal-thickness regions 101 with overlapping boundaries is not greater than a preset difference threshold.

[0055] In some alternative embodiments, the straight skin assembly of an aircraft or similar device includes four circumferentially connected panels with flush inner surfaces, such that the inner surface of the straight skin assembly is a smooth cylindrical surface. Of course, the straight skin assembly may also include other numbers of panels, as long as all panels are interconnected to form a straight section 100 with a smooth inner surface; this embodiment does not impose specific limitations.

[0056] The following section uses an example of a straight-section skin assembly for an aircraft, which consists of four interconnected panels with flush inner surfaces, to further illustrate the specific structure of such an assembly.

[0057] Specifically, the vertical skin assembly of the aircraft includes: an upper panel 110 located above the vertical section 100 of the aircraft, a right panel 120 located to the right of the vertical section 100 of the aircraft, a left panel 130 located to the left of the vertical section 100 of the aircraft, and a lower panel 140 located below the vertical section 100 of the aircraft. Due to the symmetry requirements of the aircraft, the panel corresponding to the left side of the vertical section 100 of the aircraft is completely symmetrical with respect to the panel corresponding to the right side of the vertical section 100 of the aircraft, that is, the left panel 130 and the right panel 120 are completely symmetrical with respect to the central axis of the aircraft fuselage.

[0058] Furthermore, the upper panel 110, right panel 120, left panel 130, and lower panel 140 are all laminated structures with a constant thickness in the circumferential direction, and include multiple alternating equal-thickness regions 101 and transition regions 102 in the flight direction. These transition regions 102 all have a preset slope; the minimum load thickness of each equal-thickness region 101 is obtained based on the load of each region of the straight section 100 of the aircraft, and the thickness of each equal-thickness region 101 is obtained in conjunction with the aerodynamic performance requirements of the aircraft.

[0059] The aerodynamic requirements for the aircraft refer to the flatness requirements of the straight skin components in the circumferential direction. Furthermore, since the upper panel 110, right panel 120, left panel 130, and lower panel 140 all maintain a constant thickness in the circumferential direction, the flatness of the straight skin components in the circumferential direction is actually determined based on the contact boundaries between the panels. In other words, the aerodynamic requirements of the aircraft are the flatness requirements at the contact boundaries between the panels.

[0060] For example, the flatness requirement at the boundary where each wall panel contacts another is that the thickness difference between the two wall panels at any point on the boundary is not greater than a preset difference threshold.

[0061] Based on this, as shown in Figures 3-4 and 6, in order to facilitate those skilled in the art to understand the technical solution of this embodiment, the following exemplarily provides a method for obtaining the thickness of each equal-thickness region 101 based on the corresponding minimum load thickness and in combination with the aerodynamic performance requirements of the aircraft for equal-thickness regions A and B on the upper wall panel 110 and equal-thickness region C on the right wall panel 120.

[0062] Specifically, the minimum load thickness of equal-thickness region A is 1.6 mm, the minimum load thickness of equal-thickness region B is 1.8 mm, and the minimum load thickness of equal-thickness region C is 1.8 mm. It should be noted that the overlapping portion of equal-thickness region C with equal-thickness regions A and B at the boundary is the contact boundary between the upper wall panel 110 and the right wall panel 120. Therefore, the overlapping portion of equal-thickness region C with equal-thickness regions A and B needs to meet the aerodynamic requirements of the aircraft. Based on this, the thickness difference between equal-thickness regions C and A and B is no greater than a preset difference threshold.

[0063] If the preset difference threshold is 1mm, since the minimum load thickness difference between equal thickness region C and equal thickness regions A and B is not greater than the preset difference threshold, the thicknesses of equal thickness regions A, B, and C are the corresponding minimum load thicknesses. That is, the thickness of equal thickness region A is 1.6mm, the thickness of equal thickness region B is 1.8mm, and the thickness of equal thickness region C is 1.8mm. While meeting the aerodynamic performance requirements and load-bearing requirements of the aircraft, the thickness of the straight skin components of the aircraft is relatively small, thus achieving better flight performance.

[0064] If the preset difference threshold is 0.1mm, and the difference in minimum load thickness between equal-thickness region C and equal-thickness region A is greater than this preset difference threshold, while the difference in minimum load thickness between equal-thickness region C and equal-thickness region B is not greater than this preset difference threshold, then the minimum load thickness of equal-thickness region B is taken as the thickness of equal-thickness region B, the minimum load thickness of equal-thickness region C is taken as the thickness of equal-thickness region C, and the minimum load thickness of equal-thickness region C is taken as the thickness of equal-thickness region A. That is, the thicknesses of equal-thickness regions A, B, and C are all 1.8mm, so that the aircraft... The flatness of the straight section skin assembly is further improved, thereby enhancing the aerodynamic performance of the aircraft; alternatively, the minimum load thickness of the equal-thickness region B is taken as the thickness of the equal-thickness region B, the minimum load thickness of the equal-thickness region C is taken as the thickness of the equal-thickness region C, and the difference between the minimum load thickness of the equal-thickness region C and a preset difference threshold is taken as the thickness of the equal-thickness region A. That is, the thickness of the equal-thickness region A is 1.7 mm, the thickness of the equal-thickness region B is 1.8 mm, and the thickness of the equal-thickness region C is 1.8 mm, so as to meet the aerodynamic performance requirements of the aircraft while reducing the weight of the aircraft.

[0065] Similarly, based on the above-described method for obtaining the thickness of each equal-thickness region 101, the thickness of all equal-thickness regions 101 on each panel can be obtained to obtain the structure of the entire straight section skin assembly of the aircraft. This ensures that the thickness difference at any point on the contact boundary between the panels is not greater than a preset difference threshold. Based on this, since the thickness of each panel remains constant in the circumferential direction and the flatness requirement is met at the contact boundary between the panels, the airflow of the straight section 100 of the aircraft during flight meets the requirements for normal flight of the aircraft, avoids the generation of vortices, and thus meets the aerodynamic requirements of the aircraft.

[0066] Further, as shown in Figure 5, the transition region 102 is a region on the same wall panel where the thickness of two regions 101 of equal thickness transitions smoothly. The transition region 102 is a ring-shaped surface with a preset slope arranged circumferentially. The preset slope is the slope at which the wind resistance of the straight section 100 of the aircraft is within a preset wind resistance range, in order to meet the flight requirements of the aircraft.

[0067] Specifically, the transition region 102 can be obtained based on the thickness difference of the equal-thickness regions 101 on both sides of the transition region 102, combined with a preset slope. For example, as shown in Figures 3-5, the preset slope is 1:100. For the transition region 102 between the equal-thickness regions C and D on the right wall panel 120, the method for obtaining it can be as follows: If the thickness of equal-thickness region C is 1.8mm and the thickness of equal-thickness region D is 2.0mm, then the thickness difference between equal-thickness regions C and D is 0.2mm. Therefore, the width of the transition region 102 between equal-thickness regions C and D is 20mm, and the two side boundaries of the transition region 102 are at the same height as equal-thickness regions C and D, respectively. Based on this, the transition region 102 between equal-thickness regions C and D can be obtained.

[0068] Similarly, based on the above method for obtaining the transition region 102, the transition region 102 between all adjacent equal-thickness regions 101 on each panel is obtained to obtain the structure of the entire straight skin assembly of the aircraft, thereby reducing the wind resistance of the aircraft during flight and helping the aircraft achieve better flight performance.

[0069] Based on this, the straight-section skin assembly for aircraft provided in this embodiment uses the flat side of the composite material as the inner surface of the panel, and the panels are connected to each other with their inner surfaces flush to form the straight-section skin assembly. This makes the inner surface of the straight-section skin assembly a smooth cylindrical surface, simplifying the configuration of the stringers and circumferential frames that fit the inner surface of the skin, thereby reducing the manufacturing and assembly difficulty of the stringers and circumferential frames, and helping to reduce the manufacturing cost of the aircraft. At the same time, this embodiment forms multiple alternating equal-thickness regions and transition regions on the outer surface of each panel in the flight direction. That is, the thickness of each panel does not change in the circumferential direction, but only in the flight direction. This adapts to the load requirements of different areas of the aircraft while reducing the impact of unevenness of the outer surface of the panel on the aerodynamic performance of the aircraft, thereby achieving better flight performance and benefiting the practical application of the aircraft.

[0070] Furthermore, this application also provides a vertical frame for an aircraft, including a stringer, a circumferential frame, and a skin assembly for the vertical portion of the aircraft covering the outside of the stringer and the circumferential frame. The stringer is a structure arranged along the heading of the vertical portion 100 of the aircraft, and the circumferential frame is a structure arranged circumferentially along the vertical portion 100 of the aircraft, forming the skeleton of the entire vertical portion 100 of the aircraft, thereby providing support for the skin assembly. Specifically, all stringers and all circumferential frames are fitted to the inner surface of the skin assembly of the vertical portion of the aircraft to better support it.

[0071] It should be noted that the aircraft straight-section skin assembly in this embodiment is the same as described above. Since the inner surface of the aircraft straight-section skin assembly forms a smooth cylindrical surface, the stringer and circumferential frame in this embodiment do not require complex surface designs to fit the inner surface of the aircraft straight-section skin assembly. This results in a simple configuration, fewer parts, and easier installation, effectively reducing the manufacturing difficulty and cost of the aircraft. Specifically, the stringer is a structure extending along the flight direction of the aircraft straight section 100, and the circumferential frame is a ring-shaped structure that is closed along the circumferential direction of the aircraft straight section 100.

[0072] In some optional embodiments, the straight frame of an aircraft also includes a fairing for improving the smoothness of the outer surface of the straight section 100, which at least covers the connection between the skin assembly of the straight section and the external structure. It should be noted that since the locations connecting external components often have complex structures, they have a significant impact on the aerodynamic performance of the aircraft. Therefore, this embodiment uses a fairing located on the outside of the skin assembly of the straight section to cover these complex structures, thereby reducing the impact on the aerodynamic performance of the aircraft and achieving better flight performance. For example, when the aircraft is an airplane, the fairing at least covers the connection between the aircraft fuselage skin assembly and the wing. Specifically, as shown in Figures 3-4, the fairing is generally located at the position of the equal-thickness region D on the right panel 120 and symmetrically positioned on the left panel 130.

[0073] In some alternative embodiments, the straight frame of the aircraft also includes a sealant that covers the contact boundaries of the panels of the straight skin assembly of the aircraft to fill and seal the contact boundaries of the panels, so that the thickness difference at the boundaries of the panels of the straight skin assembly of the aircraft changes smoothly, further improving the smoothness of the outer surface of the straight part 100 of the aircraft in the circumferential direction, thereby enabling the aircraft to obtain better aerodynamic performance.

[0074] Based on the straight frame of the aircraft provided in this embodiment, the main body of the aircraft is obtained. For example, an aircraft main body is obtained, as shown in Figure 7, which illustrates the airflow velocity distribution of the aircraft main body at a speed of 500-800 km / h. As shown in Figure 7, the maximum gas velocity difference between any two points on the straight frame of the aircraft provided in this embodiment, i.e., on the fuselage frame, is 30 m / s. Therefore, in this embodiment, the airflow velocity distribution on the fuselage frame is relatively uniform, meaning the pressure distribution on the fuselage surface is relatively uniform. Consequently, the aerodynamic performance of this fuselage frame is good. Furthermore, because the fuselage frame provided in this embodiment has low manufacturing costs and low assembly difficulty of the stringers and circumferential frames, the manufacturing cost of the aircraft is effectively reduced, thus benefiting the actual production and manufacturing of the aircraft.

[0075] In summary, the aircraft straight-section skin assembly and frame provided in this application, through multiple interconnected wall panels with flush inner surfaces, and because the inner surfaces of each wall panel are smooth, and the outer surfaces include multiple equally thick regions and transition regions arranged alternately along the flight direction, ensure good aerodynamic performance of the aircraft while maintaining a smooth cylindrical inner surface. This simplifies the configuration of the stringers and circumferential frames, reduces the manufacturing cost and assembly difficulty of the stringers and circumferential frames, and effectively lowers the overall manufacturing cost of the aircraft. Furthermore, this application further improves the aerodynamic performance of the aircraft through fairings and sealants, resulting in better flight performance and facilitating practical production applications.

[0076] The descriptions of the processes or structures corresponding to the above figures each have their own emphasis. For parts of a process or structure that are not described in detail, please refer to the relevant descriptions of other processes or structures.

[0077] The above embodiments are merely illustrative of the principles and effects of this application and are not intended to limit this application. Any person skilled in the art can modify or alter the above embodiments without departing from the spirit and scope of this application. Therefore, all equivalent modifications or alterations made by those skilled in the art without departing from the spirit and technical concept disclosed in this application should still be covered by the claims of this application.

Claims

1. A straight-section skin assembly for aircraft, etc., characterized in that, The system comprises multiple wall panels that are sequentially spliced ​​in a circumferential direction and have flush inner surfaces. The inner surfaces of all the wall panels form a smooth cylindrical surface. The outer surfaces of each wall panel include multiple alternating equal-thickness regions and transition regions in the flight direction. The thickness of any equal-thickness region is not less than the corresponding minimum load thickness. The surface of any transition region is a toroidal surface arranged in a circumferential direction. The two side boundaries of any transition region are in contact with the adjacent equal-thickness regions, and the thickness at the two side boundaries of the transition region is equal to the thickness of the equal-thickness regions it contacts. Each transition region has a preset slope in the flight direction. The minimum load thickness is the minimum thickness of the equal-thickness region that can support the maximum load at the corresponding position.

2. The aircraft straight-section skin assembly according to claim 1, characterized in that, The preset slope is the slope at which the wind resistance of the straight section of the aircraft is within a preset wind resistance range.

3. The aircraft straight-section skin assembly according to claim 2, characterized in that, The preset slope is 1:

100.

4. The aircraft straight-section skin assembly according to claim 1, characterized in that, The thickness difference between any two equal-thickness regions with overlapping boundaries is no greater than a preset difference threshold.

5. The aircraft straight-section skin assembly according to claim 4, characterized in that, The preset difference threshold is 1 mm.

6. The aircraft straight-section skin assembly according to claim 4, characterized in that, If the difference in minimum load thickness between any two equal-thickness regions with overlapping boundaries is not greater than the preset difference threshold, then the thickness of the equal-thickness region is the corresponding minimum load thickness.

7. The aircraft straight-section skin assembly according to claim 1, characterized in that, It includes four interlocking wall panels with flush inner surfaces, wherein the wall panel corresponding to the left side of the straight part of the aircraft is mirror-symmetrical to the wall panel corresponding to the right side of the straight part of the aircraft.

8. A straight frame for an aircraft, characterized in that, It includes the aircraft straight skin assembly as described in claims 1-7, as well as a plurality of stringers and a plurality of circumferential frames, wherein the stringers are structures extending along the flight direction of the aircraft straight section, and the circumferential frames are annular structures arranged in a closed ring along the circumferential direction of the aircraft straight section, and all the stringers and all the circumferential frames are disposed in contact with the inner surface of the aircraft straight skin assembly.

9. The straight frame for aircraft, etc., according to claim 8, characterized in that, It also includes a fairing for improving the smoothness of the outer surface of a straight section of an aircraft, the fairing covering at least the connection between the straight skin assembly of the aircraft and the external structure.

10. The straight frame for aircraft, etc., according to claim 8, characterized in that, It also includes a sealant. The straight skin assembly of the aircraft includes multiple interlocking panels. The sealant covers the contact boundary on the outer surface of adjacent panels to improve the smoothness of the outer surface of the straight part of the aircraft.

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