Aircraft equal-straight-part skin assembly and aircraft equal-straight-part frame

Abstract

The invention provides an aircraft equal-straight-part skin assembly and an aircraft equal-straight-part framework, the aircraft equal-straight-part skin assembly comprises a plurality of wall plates which are annularly and sequentially spliced, the inner surfaces of the wall plates are flush, and the inner surfaces of all the wall plates form a smooth cylindrical surface; the outer surface of each wall plate comprises a plurality of equal-thickness areas and transition areas which are alternately arranged in the heading direction, the thickness of any equal-thickness area is not smaller than the corresponding minimum load thickness, the surface of any transition area is an annular face arranged in the annular direction, and the boundaries of the two sides of any transition area make contact with the adjacent equal-thickness areas correspondingly. The thicknesses of the boundaries on the two sides of the transition areas are equal to the thicknesses of the equal-thickness areas in contact with the transition areas, and any transition area has a preset gradient in the course; therefore, under the condition that the aerodynamic performance of the aircraft is guaranteed, the configuration of the stringer and the annular frame tightly attached to the inner surface of the aircraft skin is simplified, and the manufacturing cost of the aircraft is reduced.

Description

Technical Field

[0001] The present application belongs to the field of aviation manufacturing technology, and relates to an aircraft skin structure, and in particular to a straight skin component of an aircraft and a straight frame of an aircraft. Background Art

[0002] Composite materials refer to materials composed of multiple materials with different properties or different organizational forms, and are commonly used materials for aircraft skins. In order to improve the structural efficiency of aircraft skins, that is, to reduce weight as much as possible while ensuring the strength and rigidity of aircraft skins, thereby reducing costs and improving performance, it is necessary to drop layers in the laminated structure of aircraft skins, that is, the aircraft skins have different numbers of layers at different locations, so that the thickness of each area of ​​the aircraft skin is uneven.

[0003] In the prior art, in order to ensure the aerodynamic performance of the aircraft, the aircraft skin is layered inward, that is, the number of composite material layers is selectively reduced on the inner side of the aircraft skin, so that the outer surface of the aircraft skin is a smooth surface, the airflow velocity on the aircraft surface is evenly distributed, and vortex formation is avoided. However, the structural design of this aircraft skin also makes the inner surface of the aircraft skin uneven, which in turn leads to a more complex configuration of the long stringers and the annular frame close to the inner side of the aircraft skin, and a higher manufacturing and assembly difficulty, resulting in a higher manufacturing cost of the aircraft.

[0004] Therefore, how to design an aircraft skin that simplifies the configuration of the long stringer and the annular frame while ensuring the aerodynamic performance of the aircraft is an urgent problem to be solved by technical personnel in this field. Summary of the invention

[0005] The purpose of the present application is to provide a straight skin assembly and a straight frame for aircraft, etc., for solving the problem that the existing aircraft skin has an uneven inner surface due to the inward layer-dropping structural design, which makes the configuration of the aircraft's long stringers and annular frames complex, and has high manufacturing and assembly difficulties, thereby leading to a high manufacturing cost of the aircraft.

[0006] In a first aspect, the present application provides a straight skin assembly of an aircraft, etc., comprising a plurality of wall panels which are sequentially spliced ​​in a circumferential direction and whose inner surfaces are flush, the inner surfaces of all the wall panels forming a smooth cylindrical surface, the outer surface of each wall panel including a plurality of alternatingly arranged equal thickness regions and transition regions in the heading direction, the thickness of any of the equal thickness regions not being less than the corresponding minimum load thickness, the surface of any of the transition regions being an annular surface arranged in the circumferential direction, the two side boundaries of any of the transition regions respectively contacting the adjacent equal thickness regions, and the thickness at the two side boundaries of the transition region respectively being equal to the thickness of the equal thickness regions in contact therewith, and any of the transition regions having a preset slope in the heading direction;

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

[0008] In the present application, a smooth cylindrical surface is formed through the inner surfaces of all the wall panels, which does not affect the aerodynamic performance of the aircraft. At the same time, the long stringers and annular frames that are close to the inner surfaces of the straight skin components of the aircraft have a simple configuration, fewer parts, and simple assembly, thereby reducing the manufacturing cost and assembly difficulty of the aircraft and facilitating the actual production and manufacturing of the aircraft.

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

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

[0011] In one embodiment of the present application, the thickness difference between any two equal-thickness regions with overlapping portions on their boundaries is not greater than a preset difference threshold.

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

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

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

[0015] On the second aspect, the present application also provides a straight portion frame of an aircraft, etc., including the straight portion skin assembly of the aircraft, etc. as described above, and a plurality of long spars and a plurality of annular frames, wherein the long spars are structures extending along the heading direction of the straight portion of the aircraft, etc., and the annular frames are annular structures closed and arranged in annular directions along the straight portion of the aircraft, etc., and all the long spars and all the annular frames are arranged to fit the inner surface of the straight portion skin assembly of the aircraft, etc.

[0016] In one embodiment of the present application, a fairing for improving the smoothness of the outer surface of a straight portion of an aircraft or the like is also included, and the fairing at least covers the connection between the straight portion skin assembly of the aircraft or the like and the external structure.

[0017] In one embodiment of the present application, a sealant is also included. The straight skin assembly of the aircraft, etc. includes a plurality of wall panels spliced ​​together. The sealant covers the contact boundaries on the outer surfaces of adjacent wall panels to improve the smoothness of the outer surface of the straight portion of the aircraft, etc.

[0018] As described above, the present application provides a straight skin assembly for aircraft and a straight frame for aircraft, etc., through a plurality of wall panels which are interconnected and have flush inner surfaces, so that the inner surface of the straight skin assembly for aircraft and the like is a smooth cylinder, thereby simplifying the configuration of the long stringers and the annular frames, effectively reducing the manufacturing cost and assembly difficulty of the long stringers and the annular frames, and thus reducing the production cost of the aircraft, and through the outer surface of each wall panel having multiple equal thickness areas and transition areas alternately arranged along the heading direction, while meeting the load requirements of the aircraft in different areas, the thickness of each wall panel in the annular direction is constant, so that the aircraft has good aerodynamic performance, to ensure that the inner surface of the straight skin assembly for aircraft and the like is smooth, while reducing the production cost of the aircraft, the flight effect of the aircraft is better, which is beneficial to the actual production application of the aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0019] Figure 1 Shown is a schematic diagram of a composite material layer loss structure described in an embodiment of the present application.

[0020] Figure 2 Shown is a schematic diagram of the main structure of an aircraft described in an embodiment of the present application.

[0021] Figure 3 Shown is a schematic diagram of a straight structure of an aircraft, etc. described in an embodiment of the present application.

[0022] Figure 4 Shown is a schematic diagram of an explosion of a straight structure such as an aircraft described in an embodiment of the present application.

[0023] Figure 5 Shown is a schematic diagram of an aircraft wall panel structure described in an embodiment of the present application.

[0024] Figure 6 Shown is a schematic diagram of another aircraft wall panel structure described in an embodiment of the present application.

[0025] Figure 7 Shown is a schematic diagram of airflow velocity distribution of an aircraft body described in an embodiment of the present application.

[0026] Component number description

[0027] 100 Aircraft and other straight parts

[0028] 110 Upper wall plate

[0029] 120 Right wall panel

[0030] 130 Left wall panel

[0031] 140 Lower wall panel

[0032] 101 Isothickness area

[0033] 102 Transition Zone

[0034] 200 Aircraft Head

[0035] 300 Aircraft tail DETAILED DESCRIPTION

[0036] The following describes the embodiments of the present application through specific examples, and those skilled in the art can easily understand other advantages and effects of the present application from the contents disclosed in this specification. The present application can also be implemented or applied through other different specific embodiments, and the details in this specification can also be modified or changed in various ways based on different viewpoints and applications without departing from the spirit of the present application. It should be noted that the following embodiments and features in the embodiments can be combined with each other without conflict.

[0037] It should be noted that the illustrations provided in the following embodiments are only schematic illustrations of the basic concept of the present application, and thus the drawings only show components related to the present application rather than being drawn according to the number, shape and size of components in actual implementation. In actual implementation, the type, quantity and proportion of each component may be changed at will, and the component layout may also be more complicated.

[0038] In order to improve the structural efficiency of the aircraft skin, it is necessary to remove layers from the material of the aircraft skin, that is, the laminated structure of the composite material. Specifically, Figure 1 As shown, due to the different loads at different positions of the aircraft, the required thickness of the aircraft skin is different. By reducing the number of composite material layers in the area with smaller loads, the weight of the aircraft can be effectively reduced, thereby reducing costs and improving the performance of the aircraft. However, this design will cause the surface of one side of the composite material to have undulations, thereby causing the aircraft skin to be uneven. In the prior art, in order to prevent the unevenness of the aircraft shell from affecting the gas flow velocity distribution on the aircraft surface, thereby affecting the aerodynamic performance of the aircraft, the uneven surface is generally used as the inner surface of the aircraft skin, that is, the outer surface of the aircraft skin is made into a smooth curved surface. However, this aircraft skin structure needs to match the long stringers and annular frames with complex configurations so that the long stringers and the annular frames can be close to the aircraft skin, which makes the long stringers and the annular frames difficult to manufacture and difficult to assemble, thereby resulting in a high manufacturing cost of the aircraft.

[0039] 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 interior space of the aircraft, and the outer surface of the aircraft skin is the surface of the aircraft skin facing the exterior space of the aircraft.

[0040] The following embodiments of the present application provide a straight skin component of an aircraft, etc. and a straight frame of an aircraft, etc., by forming a smooth cylinder using the flat side of a composite material as the inner surface of a straight skin component of an aircraft, etc., and the transition area of ​​the outer surface has a preset slope, so that while ensuring the aerodynamic performance of the aircraft, the long stringers and the annular frames do not need to have a complex configuration to fit with the aircraft skin, thereby simplifying the manufacture and assembly of the long stringers and the annular frames of the aircraft, effectively reducing the manufacturing cost of the aircraft, and solving the problem of high manufacturing difficulty and assembly difficulty of the long stringers and the annular frames of the aircraft in the prior art.

[0041] The following will describe in detail the principles and implementation methods of a straight skin assembly for aircraft and a straight frame for aircraft of this embodiment in conjunction with the accompanying drawings, so that those skilled in the art can understand the straight skin assembly for aircraft and the straight frame for aircraft of this embodiment without creative work.

[0042] like Figure 1 As shown, 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 has an uneven surface.

[0043] like Figure 2 It is a structural schematic diagram of the aircraft body, which includes: an aircraft head 200, an aircraft straight portion 100 and an aircraft tail 300. The aircraft head 200 is located at the front of the aircraft, and the aircraft tail 300 is located at the tail of the aircraft. The aircraft head 200 and the aircraft tail 300 are connected through the aircraft straight portion 100. Among them, the aircraft head 200, the aircraft straight portion 100 and the aircraft tail 300 are all made of 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 long stringer and the annular frame that fit the inner surface of the aircraft skin assembly, so as to reduce the manufacturing difficulty and assembly difficulty of the long stringer and the annular frame, thereby reducing the manufacturing cost of the aircraft.

[0044] It should be noted that, in order to reduce the flight resistance of the aircraft, the shape of the aircraft body is usually a streamlined shape with less flight resistance. Specifically, the curvature of each point on the aircraft head 200 changes smoothly, and the aircraft tail 300 is connected by a columnar aircraft straight portion 100. For example, Figure 2As shown, when the aircraft is an airplane, the curvature of each point on the aircraft head 200 and the aircraft tail 300 changes smoothly, and the surface of the aircraft straight portion 100 is a cylinder, and the curvature of each point remains unchanged. 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, wherein the portion with a large curvature change is more sensitive to airflow changes, while the area with a constant curvature has a smaller impact on the aerodynamic performance, that is, the flatness requirements of the outer surface of the aircraft straight portion 100 are much lower than those of the aircraft head 200 and the aircraft tail 300.

[0045] In the present embodiment, a straight portion skin assembly of an aircraft or the like located at a straight portion 100 of the aircraft or the like is taken as an example to explain its specific structure and working principle.

[0046] like Figure 3-4 As shown, the straight skin assembly of the aircraft or the like includes a plurality of wall panels connected to each other and having flush inner surfaces, and the inner surfaces of these wall panels form a smooth cylindrical surface to simplify the configuration of the long stringer and the configuration of the annular frame that fit the inner surface. It should be noted that each wall panel can be manufactured in an integral molding manner, and the entire straight skin assembly of the aircraft or the like can be obtained by splicing each wall panel.

[0047] In order to reduce the weight of the aircraft, the laminated structure of the composite material needs to reduce the weight by dropping layers, so the outer surfaces of these wall panels are actually uneven surfaces. Specifically, the outer surface of each wall panel includes a plurality of alternating equal thickness regions 101 and transition regions 102 in the heading direction, wherein the equal thickness region 101 is a region with equal thickness everywhere, and the surface of the transition region 102 is a ring surface arranged along the circumferential direction. Based on this, the outer surface of a single wall panel has no thickness change in the circumferential direction, but only changes the thickness in the heading direction. It should be noted that when there are thickness changes in multiple directions on the outer surface of the skin, the flow of airflow is more complicated, which will have a greater impact on the aerodynamic performance of the aircraft, and even form vortices on the surface of the aircraft, affecting the normal flight of the aircraft. Based on this, in this embodiment, the outer surface of each wall panel of the aircraft has a thickness change only in the heading direction, and the thickness remains unchanged in the circumferential direction, thereby avoiding the thickness changes in multiple directions on the outer surface of the aircraft, thereby reducing the impact of the unevenness of the outer surface of the aircraft on the aerodynamic performance of the aircraft, that is, the impact on the aerodynamic performance of the aircraft is minimal. Based on this, the thickness of each wall panel in the present embodiment does not change in the circumferential direction, that is, the thickness of each wall panel in the circumferential direction is constant, and the thickness only changes in the heading direction, so that the thickness of each wall panel changes based on the load to ensure the structural efficiency of each area while avoiding affecting the aerodynamic performance of the aircraft, thereby obtaining a better flight effect. Moreover, since the inner surface of the aircraft skin is smooth and the long stringers and the circumferential frame have simple configurations, the manufacturing cost of the aircraft is effectively reduced, which is beneficial to the practical application of the aircraft.

[0048] The heading is the direction in which the aircraft is flying, and the circumferential direction is the circumferential direction around the central axis of the aircraft fuselage.

[0049] 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 heading direction, the boundaries of the transition region 102 on both sides are in contact with the adjacent equal thickness regions 101, and the thickness at the boundaries of the transition region on both sides is equal to the thickness of the corresponding equal thickness region 101, so as to achieve a smooth transition of each equal thickness region 101. Among them, 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 the corresponding position on the aircraft bears the maximum load is the minimum load thickness of the corresponding equal thickness region 101.

[0050] It should be noted that if Figure 1 As shown, the transition region 102 is a region formed due to layer loss of the laminated structure of the aircraft skin, and the width of the transition region 102 is much smaller than the equal thickness region 101. Figure 5 As shown, the transition area 102 connects the boundaries of two equal-thickness areas 101. If the boundary thickness difference between the two equal-thickness areas 102 is h1, that is, the thickness difference between the two side boundaries of the transition area 102 is h1, and the width of the transition area 102 is x, then the annular slope of the transition area is the ratio of h1 to x, that is, the preset slope is h1:x.

[0051] It should be noted that, in this embodiment, the preset slope is calculated based on the thickness difference and width between the two side boundaries of the transition region 102. However, in the actual preparation 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 side boundaries of the transition region 102, and based on the preset slope, the width of the transition region 102 is obtained.

[0052] Preferably, the preset slope is a slope at which the wind resistance of the straight portion 100 of the aircraft during the flight of the aircraft is within a preset wind resistance range, wherein 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, and those skilled in the art should be aware of it, and will not be elaborated here.

[0053] It should be noted that, since the annular surface of the transition region 102 will form a certain wind resistance during the flight, affecting the flight of the aircraft, in this embodiment, the slope of the annular surface of the transition region 102 is set to a slope that meets the requirements of smooth airflow, so as to reduce the wind resistance of the aircraft during the flight, thereby further enhancing the aerodynamic performance of the aircraft.

[0054] For example, the preset slope is 1:100, that is, Figure 5As shown, h1:x is 1:100. If the thickness of the equal-thickness regions 101 on both sides changes by 1 mm, the width of the transition region 102 is 100 mm. At this time, the thickness change transition between the equal-thickness regions 101 is relatively smooth, and the wind resistance generated during the flight of the aircraft is very small, which basically does not affect the normal flight of the aircraft, thereby achieving a better flight effect.

[0055] In some optional embodiments, since the aircraft skin includes multiple wall panels and different areas of the aircraft have load requirements, the thickness of each wall panel is different and the contacting wall panel boundaries have a thickness difference. In order to further simplify the configuration of the long stringers and the annular frames, in this embodiment, each is avoided from being connected to each other and the inner surfaces are flush, so that the inner surface of the straight skin component of the aircraft forms a smooth cylinder, thereby making the long stringers and the annular frames that fit the inner surface of the straight skin component of the aircraft simple in structure, effectively reducing the manufacturing cost of the aircraft.

[0056] It should be noted that if Figure 6 As shown, since there may be thickness differences at the boundaries where the wall panels contact each other, the outer surface of the aircraft skin is uneven, wherein the thickness difference at any point on the boundaries where the wall panels contact each other is not greater than a preset difference threshold, so as to minimize the difference in outer surface flatness of the straight portion 100 of the aircraft caused by the thickness difference between the wall panels, thereby further improving the aerodynamic performance of the aircraft.

[0057] It should be noted that if Figure 6 As shown, the thickness difference on the boundary where each wall panel contacts each other is actually: the thickness difference of the equal thickness area with overlapping parts on the boundary of two adjacent wall panels, that is, the value of h2. Based on this, in order to ensure that the thickness difference at any point on the boundary where each wall panel contacts each other is not greater than the preset difference threshold, in some optional embodiments, for any two equal thickness areas 101 with overlapping parts on the boundary, the thickness difference between the two is not greater than the preset difference threshold, so as to achieve that the thickness difference at any point on the boundary where each wall panel contacts each other is not greater than the preset difference threshold.

[0058] It should be noted that, since the equal-thickness regions 101 and the transition regions 102 on the same wall panel are arranged alternately, there is no overlapping boundary between any two equal-thickness regions 101 on the same wall panel, and the equal-thickness regions 101 with overlapping boundaries are actually located on two adjacent wall panels. Figure 3-4 As shown, the boundaries of the equal-thickness region C and the equal-thickness region A have an overlapping portion, and similarly, the boundaries of the equal-thickness region C and the equal-thickness region B have an overlapping portion.

[0059] Exemplarily, the preset difference threshold is 1 mm, that is, for any two equal-thickness regions 101 with overlapping boundaries, the thickness difference does not exceed 1 mm.

[0060] It should be noted that if the difference in the minimum load thickness between any two equal-thickness regions 101 with overlapping borders is not greater than the preset difference threshold, the thickness of the equal-thickness region 101 is the corresponding minimum load thickness. That is, the minimum load thickness of all equal-thickness regions 101 is calculated based on the loads of each region of the equal straight portion 100 of the aircraft, and the difference in the minimum load thickness between all equal-thickness regions 101 with overlapping borders is obtained. If these differences are not greater than the preset difference threshold, the thickness of all equal-thickness regions 101 is the corresponding minimum load thickness, so as to reduce the thickness of all equal-thickness regions 101, thereby reducing the weight of the aircraft, which is conducive to the aircraft obtaining a better flight effect.

[0061] 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.

[0062] Of course, the maximum value of all minimum load thicknesses can also be obtained, and the maximum value can be used as the thickness of the corresponding equal thickness region 101, and the thickness of other equal thickness regions 101 can be obtained based on the preset difference threshold. For example, the difference between the maximum value and the preset difference threshold is used as the thickness of the equal thickness region 101 that overlaps or partially overlaps with the boundary of the equal thickness region 101 corresponding to the maximum value, and so on to obtain the thickness of all other equal thickness regions 101. Based on this, the weight of the aircraft can be reduced while ensuring the aerodynamic performance of the aircraft, so that the aircraft can obtain a good flight effect.

[0063] It should be noted that the above embodiment exemplifies a method for obtaining the thickness of the equal-thickness region 101, but is not limited thereto. A person skilled in the art may 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.

[0064] In some optional embodiments, the aircraft and other straight skin components include four wall panels that are sequentially spliced ​​in a circumferential direction and have flush inner surfaces, so that the inner surface of the aircraft and other straight skin components is a smooth cylinder. Of course, the aircraft and other straight skin components may also include other numbers of wall panels, as long as all the wall panels are connected to each other to form an aircraft and other straight portion 100 with a smooth inner surface, and this embodiment does not make any specific restrictions here.

[0065] The following takes an example in which a straight skin assembly of an aircraft includes four wall panels that are interconnected and have flush inner surfaces to further explain the specific structure of the straight skin assembly of an aircraft.

[0066] Specifically, the aircraft straight portion skin assembly includes: an upper wall panel 110 located in the area above the aircraft straight portion 100, a right wall panel 120 located in the area to the right of the aircraft straight portion 100, a left wall panel 130 located in the area to the left of the aircraft straight portion 100, and a lower wall panel 140 located in the area below the aircraft straight portion 100, wherein, due to the symmetry requirement of the aircraft, the wall panel corresponding to the left side of the aircraft straight portion 100 is completely symmetrical with the wall panel corresponding to the right side of the aircraft straight portion 100, that is, the left wall panel 130 is completely symmetrical with the right wall panel 120 relative to the central axis of the aircraft fuselage.

[0067] Furthermore, the upper wall plate 110, the right wall plate 120, the left wall plate 130 and the lower wall plate 140 are all laminated structures with constant thickness in the circumferential direction, and include a plurality of alternating equal thickness regions 101 and transition regions 102 in the heading direction. Among them, these transition regions 102 all have a preset slope; based on the load of each region of the equal straight portion 100 of the aircraft, the minimum load thickness of each equal thickness region 101 is obtained, and combined with the aerodynamic performance requirements of the aircraft, the thickness of each equal thickness region 101 is obtained.

[0068] Among them, the aerodynamic requirements of the aircraft are the flatness requirements of the straight skin components of the aircraft in the circumferential direction. Furthermore, since the thickness of the upper wall panel 110, the right wall panel 120, the left wall panel 130 and the lower wall panel 140 in the circumferential direction remains unchanged, the flatness of the straight skin components of the aircraft in the circumferential direction is actually determined based on the contact boundaries between the wall panels, that is, the aerodynamic requirements of the aircraft are the flatness requirements at the contact boundaries between the wall panels.

[0069] Exemplarily, the flatness requirement at the boundary where the wall panels contact each other is as follows: at any point on the boundary, the thickness difference between the wall panels on both sides is not greater than a preset difference threshold.

[0070] Based on this, Figure 3-4 as well as Figure 6 As shown, in order to facilitate those skilled in the art to understand the technical solution of the present embodiment, the following exemplarily provides a method for obtaining the thickness of each equal thickness area 101 based on the corresponding minimum load thickness and in combination with the aerodynamic performance requirements of the aircraft for equal thickness areas A and B on the upper wall panel 110, and equal thickness area C on the right wall panel 120.

[0071] Specifically, the minimum load thickness of the equal thickness region A is 1.6 mm, the minimum load thickness of the equal thickness region B is 1.8 mm, and the minimum load thickness of the equal thickness region C is 1.8 mm. It should be noted that the overlapping portion of the equal thickness region C on the boundary with the equal thickness region A and the equal thickness region B is the portion of the boundary where the upper wall panel 110 and the right wall panel 120 contact. Therefore, the overlapping portion of the equal thickness region C on the boundary with the equal thickness region A and the equal thickness region B needs to meet the aerodynamic requirements of the aircraft. Based on this, the thickness difference between the equal thickness region C and the equal thickness region A and the equal thickness region B is not greater than the preset difference threshold.

[0072] If the preset difference threshold is 1mm, since the minimum load thickness differences between the equal-thickness region C and the equal-thickness region A and the equal-thickness region B are no more than the preset difference threshold, the thicknesses of the equal-thickness region A, the equal-thickness region B and the equal-thickness region C are the corresponding minimum load thicknesses, that is, the thickness of the equal-thickness region A is 1.6mm, the thickness of the equal-thickness region B is 1.8mm, and the thickness of the equal-thickness region C is 1.8mm. While meeting the aerodynamic performance requirements and the load-bearing requirements of the aircraft, the thickness of the straight skin components of the aircraft is relatively small, thereby achieving a better flight effect.

[0073] If the preset difference threshold is 0.1mm, the difference in the minimum load thickness between the equal thickness region C and the equal thickness region A is greater than the preset difference threshold, and the difference in the minimum load thickness between the equal thickness region C and the equal thickness region B is not greater than the preset difference threshold, then the minimum load thickness of the equal thickness region B is used as the thickness of the equal thickness region B, the minimum load thickness of the equal thickness region C is used as the thickness of the equal thickness region C, and the minimum load thickness of the equal thickness region C is used as the thickness of the equal thickness region A, that is, the thicknesses of the equal thickness regions A, B, and C are all 1.8mm, so that the aircraft The flatness of the equal straight skin assembly is further improved, thereby improving the aerodynamic performance of the aircraft; alternatively, the minimum load thickness of the equal thickness region B is used as the thickness of the equal thickness region B, the minimum load thickness of the equal thickness region C is used as the thickness of the equal thickness region C, and the difference between the minimum load thickness of the equal thickness region C and the preset difference threshold is used 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 and reduce the weight of the aircraft.

[0074] Similarly, based on the above-mentioned method for obtaining the thickness of each equal-thickness area 101, the thickness of all equal-thickness areas 101 on each wall panel can be obtained to obtain the structure of the entire aircraft straight skin assembly, so that the thickness difference at any point on the contact boundary between each wall panel is not greater than the preset difference threshold. Based on this, since the thickness of each wall panel remains unchanged in the circumferential direction and the contact boundary between each wall panel meets the flatness requirement, the flow condition of the airflow of the aircraft straight portion 100 during flight meets the requirements of normal flight of the aircraft and avoids the generation of vortices, that is, it meets the aerodynamic requirements of the aircraft.

[0075] Furthermore, if Figure 5 As shown, the transition area 102 is an area on the same wall panel where the thickness smoothly transitions between two equal-thickness areas 101. The transition area 102 is an annular surface with a preset slope and arranged along the annular direction, and the preset slope is a slope where the wind resistance of the straight portion 100 of the aircraft is within a preset wind resistance range to meet the flight requirements of the aircraft.

[0076] Specifically, based on the thickness difference of the equal-thickness regions 101 on both sides of the transition region 102 and the preset slope, the transition region 102 can be obtained. Figure 3-5 As shown, the preset slope is 1:100. For the transition area 102 between the equal thickness area C and the equal thickness area D on the right wall plate 120, the acquisition method can be: if the thickness of the equal thickness area C is 1.8 mm, the thickness of the equal thickness area D is 2.0 mm, and the thickness difference between the equal thickness area C and the equal thickness area D is 0.2 mm, then the width of the transition area 102 between the equal thickness area C and the equal thickness area D is 20 mm, and the two side boundaries of the transition area 102 are respectively at the same height as the equal thickness area C and the equal thickness area D. Based on this, the transition area 102 between the equal thickness area C and the equal thickness area D can be obtained.

[0077] Similarly, based on the above-mentioned method of obtaining the transition area 102, the transition area 102 between all adjacent equal-thickness areas 101 on each wall panel is obtained to obtain the structure of the straight skin components of the entire aircraft, thereby reducing the wind resistance of the aircraft during flight, which is conducive to the aircraft achieving better flight effects.

[0078] Based on this, the straight skin assembly of aircraft and the like provided in this embodiment uses the flat side of the composite material as the inner surface of the wall panel, and each wall panel is connected to each other in a manner that the inner surface is flush to form a straight skin assembly of aircraft and the like, so that the inner surface of the straight skin assembly of aircraft and the like is a smooth cylinder, which simplifies the configuration of the long stringer and the annular frame that fits the inner surface of the skin, thereby reducing the manufacturing and assembly difficulties of the long stringer and the annular frame, which is conducive to reducing the manufacturing cost of the aircraft. At the same time, this embodiment forms a plurality of alternating equal thickness areas and transition areas on the outer surface of each wall panel in the heading direction, that is, the thickness of each wall panel does not change in the annular direction, but only changes in the heading direction. While adapting to the load requirements of different areas of the aircraft, it reduces the impact of the uneven outer surface of the wall panel on the aerodynamic performance of the aircraft, thereby obtaining a better flight effect, which is conducive to the practical application of the aircraft.

[0079] Furthermore, the present application also provides a straight part frame of an aircraft, etc., including a long stringer, an annular frame, and an aircraft, etc. straight part skin assembly covering the outside of the long stringer and the annular frame. Among them, the long stringer is a structure arranged along the heading direction of the straight part 100 of the aircraft, etc., and the annular frame is a structure arranged along the annular direction of the straight part 100 of the aircraft, etc., to form the skeleton of the entire straight part 100 of the aircraft, etc., so as to provide support for the straight part skin assembly of the aircraft, etc. Specifically, all the long stringers and all the annular frames are arranged in contact with the inner surface of the straight part skin assembly of the aircraft, etc., so as to better provide support for the straight part skin assembly of the aircraft, etc.

[0080] It should be noted that the aircraft and other straight skin components in this embodiment are the aircraft and other straight skin components described above. Since the inner surface of the aircraft and other straight skin components forms a smooth cylindrical surface, the long stringer and the annular frame in this embodiment do not need to design a complex surface to fit the inner surface of the aircraft and other straight skin components. The configuration is simple, the number of parts is small, and it is easy to install, which effectively reduces the manufacturing difficulty and cost of the aircraft. Specifically, the long stringer is a structure extending along the heading direction of the aircraft and other straight parts 100, and the annular frame is an annular structure that is closed and arranged along the annular direction of the aircraft and other straight parts 100.

[0081] In some optional embodiments, the straight portion frame of the aircraft further includes a fairing for improving the smoothness of the outer surface of the straight portion 100 of the aircraft, and the fairing at least covers the connection between the straight portion skin assembly of the aircraft and the external structure. It should be noted that since the location where the external components are connected usually has a relatively complex structure, it has a great impact on the aerodynamic performance of the aircraft. Based on this, this embodiment covers these complex structures by setting the fairing on the outside of the straight portion skin assembly of the aircraft, thereby reducing the impact on the aerodynamic performance of the aircraft and obtaining a better flight effect. Exemplarily, when the aircraft is an airplane, the fairing at least covers the connection between the aircraft fuselage skin assembly and the wing. Specifically, as Figure 3-4As shown, the fairing is generally arranged at the position of the equal thickness area D on the right wall panel 120 and at a symmetrical position on the left wall panel 130 .

[0082] In some optional embodiments, the straight portion frame of the aircraft and the like also includes a sealant, which covers the boundaries where the wall panels of the straight portion skin assembly of the aircraft and the like contact each other, so as to fill in and seal the boundaries where the wall panels contact each other, so that the thickness difference at the boundaries of the wall panels of the straight portion skin assembly of the aircraft and the like changes smoothly, further improving the flatness of the outer surface of the straight portion 100 of the aircraft and the like in the circumferential direction, thereby enabling the aircraft to obtain better aerodynamic performance.

[0083] Based on the straight frame of the aircraft provided in this embodiment, an aircraft body is obtained. For example, an aircraft body is obtained, such as Figure 7 As shown, it is a schematic diagram of the airflow velocity distribution of the aircraft body when the speed is 500-800km / h. Figure 7 It can be known that on the straight frame of the aircraft provided in this embodiment, that is, on an aircraft fuselage frame, the maximum gas flow velocity difference between any two points is 30m / s. Based on this, in this embodiment, the air flow velocity distribution on the fuselage frame is relatively uniform, that is, the pressure distribution on the fuselage surface is relatively uniform, so the aerodynamic performance of the fuselage frame is better. In addition, since the fuselage frame provided in this embodiment has a low manufacturing cost and the assembly difficulty of the long stringers and the annular frames is low, the manufacturing cost of the aircraft is effectively reduced, which is beneficial to the actual production and manufacturing of the aircraft.

[0084] In summary, the aircraft and other straight skin components and aircraft and other straight frames provided by the present application are formed by a plurality of wall panels connected to each other and having flush inner surfaces, and because the inner surfaces of each wall panel are smooth, the outer surface includes a plurality of equal thickness areas and transition areas arranged alternately along the heading direction, so that while ensuring that the aircraft has good aerodynamic performance, the inner surface of the aircraft and other straight skin components is a smooth cylinder, thereby simplifying the configuration of the long stringer and the annular frame, reducing the manufacturing cost and assembly difficulty of the long stringer and the annular frame, and thus effectively reducing the overall manufacturing cost of the aircraft. At the same time, the present application further improves the aerodynamic performance of the aircraft through the fairing and sealant, so that the aircraft obtains a better flight effect, which is conducive to the actual production and application of the aircraft.

[0085] The descriptions of the processes or structures corresponding to the above-mentioned figures have different emphases. For parts that are not described in detail in a certain process or structure, please refer to the relevant descriptions of other processes or structures.

[0086] The above embodiments are merely illustrative of the principles and effects of the present application and are not intended to limit the present application. Anyone familiar with the technology may modify or change the above embodiments without violating the spirit and scope of the present application. Therefore, all equivalent modifications or changes made by a person of ordinary skill in the art without departing from the spirit and technical ideas disclosed in the present application shall still be covered by the claims of the present application.

Claims

1. A straight skin component for an aircraft, etc., characterized in that: It comprises a plurality of wall panels which are sequentially spliced ​​in an annular direction and whose inner surfaces are flush, the inner surfaces of all the wall panels form a smooth cylindrical surface, the outer surface of each wall panel comprises a plurality of alternating equal-thickness regions and transition regions in the heading direction, the thickness of any equal-thickness region is not less than the corresponding minimum load thickness, the surface of any transition region is an annular surface arranged in the annular direction, the two side boundaries of any transition region are respectively in contact with the adjacent equal-thickness regions, and the thickness at the two side boundaries of the transition region is respectively equal to the thickness of the equal-thickness regions in contact with them, and any transition region has a preset slope in the heading direction; The minimum load thickness is the minimum thickness of the equal thickness area that can support the maximum load at the corresponding position.

2. The straight skin assembly of an aircraft or the like according to claim 1, characterized in that: The preset slope is a slope at which the wind resistance of the straight portion of the aircraft is within a preset wind resistance range.

3. The straight skin assembly for aircraft or the like according to claim 2, characterized in that: The preset slope is 1:

100.

4. The straight skin assembly for aircraft or the like according to claim 1, characterized in that: The thickness difference between any two equal-thickness regions with overlapping borders is not greater than a preset difference threshold.

5. The straight skin assembly for aircraft or the like according to claim 4, characterized in that: The preset difference threshold is 1 mm.

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

7. The straight skin assembly of an aircraft or the like according to claim 1, characterized in that: It comprises four wall panels which are connected to each other and whose inner surfaces are flush, wherein the wall panel corresponding to the left side of the straight part of the aircraft is mirror-symmetrical with the wall panel corresponding to the right side of the straight part of the aircraft.

8. A straight frame for aircraft, etc., characterized in that: It comprises a straight skin assembly of an aircraft or the like as described in claims 1-7, and a plurality of long stringers and a plurality of annular frames, wherein the long stringers are structures extending along the heading direction of the straight portion of the aircraft or the like, and the annular frames are annular structures closed and arranged in annular directions along the straight portion of the aircraft or the like, and all the long stringers and all the annular frames are arranged to fit the inner surface of the straight skin assembly of the aircraft or the like.

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

10. The straight frame of an aircraft or the like according to claim 8, characterized in that: It also includes sealant. The straight skin assembly of the aircraft includes multiple wall panels spliced ​​together. The sealant covers the contact boundaries on the outer surfaces of adjacent wall panels to improve the smoothness of the outer surface of the straight part of the aircraft.

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