Wing structure parameter design method considering covering cap bearing efficiency

Abstract

The invention belongs to the field of aircraft structure design, and particularly relates to a wing structure parameter design method considering covering cap bearing efficiency. The method comprises the following steps: step 1, acquiring initial bearing coefficients of covering caps at different positions of a wing; 2, constructing a wing structure finite element model, and correcting the Young modulus of a covering cap material in the wing structure finite element model according to the initial bearing coefficient; step 3, performing wing structure parameter optimization according to the wing structure finite element model to obtain wing structure parameters, extracting a stress level in the wing structure parameter optimization process, obtaining a new bearing coefficient of the covering cap under the stress level, comparing the new bearing coefficient with an initial bearing coefficient, if a convergence condition is not met, returning to the step 2, and if a convergence condition is not met, returning to the step 3; and correcting the Young modulus of the covering cap material in the wing structure finite element model according to the new bearing coefficient until wing structure parameters meeting convergence conditions are obtained.

Description

Technical Field

[0001] This application belongs to the field of aircraft structural design, and specifically relates to a method for designing wing structural parameters that takes into account the load-bearing efficiency of the wing cover. Background Technology

[0002] To meet maintenance requirements, aircraft wings typically require multiple openings in the fuselage panels. For small and medium-sized aircraft, due to wing size limitations and the need to minimize structural weight, a load-bearing opening design is usually adopted (the opening and the fuselage panel jointly transfer the wing bending moment). This leads to the following issues in the parameter design of the fuselage panels and openings:

[0003] 1. The connection between the cover and the wall panel is often made with a large clearance. The load-bearing cover and the wall panel participate in the bending moment transmission of the wing, and their load transmission efficiency is different. This is especially true for wings with small chord dimensions (the cover occupies a large cross-sectional area of ​​the wall panel). If the difference in load transmission efficiency between the cover and the wall panel is ignored, it will lead to unreasonable results in the optimization design of the wing wall panel basic parameters.

[0004] 2. The load-bearing efficiency of the load-bearing cover is related to its location and stress level. The load-bearing efficiency of the cover needs to be incorporated into the parameter optimization of the wing box section for different covers, so as to ensure the reliability of the parameter design results.

[0005] 3. Based on the optimized basic structural parameters of the skin and cover area, and considering design factors such as opening form, reinforcement width and stress level, it is necessary to solve the design of detailed parameters from the optimized basic parameters to the wall panel and cover area.

[0006] Existing technologies do not quantify the impact of the cover load-bearing efficiency on the overall force transmission and lack a parametric design method for wing panel structures that consider the cover load-bearing capacity, resulting in insufficient analysis accuracy.

[0007] Therefore, there is an urgent need for a technical solution to overcome or mitigate at least one of the aforementioned defects in the existing technology. Summary of the Invention

[0008] The purpose of this application is to provide a method for designing wing structural parameters that takes into account the load-bearing efficiency of the flap, so as to solve at least one problem existing in the prior art.

[0009] The technical solution of this application is:

[0010] A method for designing wing structural parameters considering the load-bearing efficiency of the flap, comprising:

[0011] Step 1: Obtain the initial load-bearing coefficient of the hatches at different locations on the wing;

[0012] Step 2: Construct a finite element model of the wing structure and correct the Young's modulus of the wing cover material in the finite element model based on the initial load-bearing coefficient;

[0013] Step 3: Optimize the wing structure parameters based on the finite element model of the wing structure to obtain the wing structure parameters. Extract the stress level during the optimization process and obtain the new load-bearing coefficient of the hatch at this stress level. Compare the new load-bearing coefficient with the initial load-bearing coefficient. If the convergence condition is not met, return to Step 2. Correct the Young's modulus of the hatch material in the finite element model of the wing structure based on the new load-bearing coefficient until the wing structure parameters that meet the convergence condition are obtained.

[0014] In a preferred embodiment of this application, step one, obtaining the initial load-bearing coefficient of the wing openings at different locations, includes:

[0015] S101. Determine the method for obtaining the bearing capacity coefficient of the cover under different stress levels;

[0016] S102. Obtain the stress level of the cover area at different locations on the wing;

[0017] S103. Determine the initial load-bearing coefficient of the cover at different positions of the wing based on the stress level.

[0018] In a preferred embodiment of this application, in S101, the load-bearing coefficient of the cap under different stress levels is obtained by using a box segment-level test, including:

[0019] Obtain box-level test specimens for box-level testing;

[0020] Box segment level tests were conducted under different stress levels to obtain box segment level test data;

[0021] The load-bearing coefficient is calculated based on the test data of the box section.

[0022] In a preferred embodiment of this application, the box segment-level test specimen includes:

[0023] A wing beam, wherein the wing beam has openings in its wall panels and a fixed end is provided at the end of the wing beam;

[0024] A first cover is installed at the opening of the wall panel;

[0025] The first strain gauge is installed on the first cover;

[0026] The first skin strain gauge is disposed on the wall panel, and both the first skin strain gauge and the first cover strain gauge are located on the axis of symmetry of the cover.

[0027] In a preferred embodiment of this application, in S101, the bearing capacity coefficient of the cap under different stress levels is obtained by using a stiffened plate level test, including:

[0028] Obtain stiffened plate-level test specimens for stiffened plate-level testing;

[0029] Stiffened plate level tests were conducted under different stress levels to obtain stiffened plate level test data;

[0030] The bearing capacity coefficient is calculated based on the test data of the stiffened slab.

[0031] In a preferred embodiment of this application, the stiffened plate-level test specimen includes:

[0032] A stiffening plate, wherein the stiffening plate has a stiffening plate opening, the stiffening plate has a dummy rib support at the rib, and the stiffening plate has a fixed end at the end;

[0033] A second cover is installed at the opening of the stiffening plate;

[0034] The second strain gauge is installed on the second cover.

[0035] The second skin strain gauge is disposed on the stiffening plate, and both the second skin strain gauge and the second cover strain gauge are parallel to the wing rib.

[0036] In a preferred embodiment of this application, the calculated stress level is:

[0037] ;

[0038] in, For skin stress level, E sk The Young's modulus of the skin material. The strain value of the skin strain gauge. This refers to the number of skin strain gauges;

[0039] Calculate the load-bearing capacity of the cover at this stress level, including:

[0040] ;

[0041] in, For bearing capacity, The strain value of the strain gauge on the lid. The number of strain gauges for the lid. For the thickness of the cap, This refers to the skin thickness.

[0042] In a preferred embodiment of this application, step S102, obtaining the stress level of the wing cover area at different locations, includes:

[0043] Calculate skin thickness:

[0044] ;

[0045] Calculate the skin stress level:

[0046] ;

[0047] in, For skin stress level, E is the critical compressive stress coefficient. sk Let M be the Young's modulus of the skin material, M be the bending moment of the airfoil section at the hatch, μ be the Poisson's ratio of the skin material, b be the width of the loaded edge of the skin, and A be the bending moment of the airfoil section at the hatch. box Let t be the cross-sectional area of ​​the wing box section at the hatch, λ be the ratio of the area of ​​the stringer to the area of ​​the skin section, and t be the cross-sectional area of ​​the wing box section at the hatch. min This refers to the minimum processing thickness of the skin.

[0048] Determine the stress level in the flap area based on the skin stress level:

[0049] ;

[0050] in, The stress level in the lid area, This is the stress level reduction factor. This represents the allowable stress of the wall panel.

[0051] In a preferred embodiment of this application, step two, correcting the Young's modulus of the wing structure finite element model based on the initial load-bearing coefficient, includes:

[0052] For the hatch at the i-th wing position, the corrected Young's modulus of the hatch material is:

[0053] ;

[0054] Among them, E i Let E be the Young's modulus of the cover material after correction at the i-th wing position. Let be the initial load-bearing coefficient of the cover at the i-th wing position.

[0055] In a preferred embodiment of this application, the convergence condition in step three includes:

[0056] ;

[0057] in, Let be the new load-bearing coefficient for the cover at the i-th wing position.

[0058] In a preferred embodiment of this application, step three includes convergence conditions such as strength, stiffness, stability, and manufacturability requirements.

[0059] In a preferred embodiment of this application, in step three, the wing structure parameters include: the thickness of the skin on both sides of the hatch, the thickness of the hatch, and the area of ​​the stringers on both sides of the hatch.

[0060] In a preferred embodiment of this application, the method further includes step five: performing reinforcement and optimization design on the cover area structure to obtain the reinforced and optimized wing structure parameters.

[0061] In a preferred embodiment of this application, the reinforced and optimized wing structure parameters are as follows:

[0062] ;

[0063] ;

[0064] ;

[0065] ;

[0066] in, , To reinforce the thickness of the skin on both sides of the cap after optimization, , The thickness of the skin on both sides of the flap, b is the skin reinforcement coefficient. p b1 and b2 are the widths of the opening and the skin reinforcement, respectively. , To compensate for the area of ​​the stringers on both sides of the optimized cover, , The area of ​​the long stringers on both sides of the lid, This is the reinforcement coefficient for the stringer.

[0067] In a preferred embodiment of this application, the reinforcement coefficient is determined as follows: ,and .

[0068] The invention has at least the following beneficial technical effects:

[0069] This application presents a wing structure parameter design method that considers the load-bearing efficiency of the cover. This method incorporates the influence of the cover's load-bearing efficiency on the overall load transfer of the wall panel during the wing structure parameter optimization design phase. By incorporating the cover's load-bearing efficiency, more reliable parameters can be obtained during the wing structure parameter optimization design stage. Simultaneously, considering issues such as stress concentration at the opening edge, the cover area structure is reinforced to increase the stiffness and strength of the opening area. Attached Figure Description

[0070] Figure 1 This is a flowchart of a wing structure parameter design method considering the load-bearing efficiency of the cover, according to one embodiment of this application.

[0071] Figure 2 This is a schematic diagram of a box-section level test specimen and a stiffening plate level test specimen according to one embodiment of this application;

[0072] Figure 3 This is a schematic diagram of the strain gauge arrangement of a box-section test specimen according to one embodiment of this application;

[0073] Figure 4 This is a schematic diagram of the strain gauge arrangement of a stiffened plate-level test specimen according to one embodiment of this application;

[0074] Figure 5 This is a schematic diagram of a finite element model of an airfoil structure according to one embodiment of this application;

[0075] Figure 6 This is a schematic diagram of wing structure parameters according to one embodiment of this application;

[0076] Figure 7 This is a schematic diagram of the wing structure parameter reinforcement design according to one embodiment of this application;

[0077] Figure 8 This is a schematic diagram showing the detailed structure of the cover area wall panel according to one embodiment of this application. Detailed Implementation

[0078] To make the objectives, technical solutions, and advantages of this application clearer, the technical solutions in the embodiments of this application will be described in more detail below with reference to the accompanying drawings. In the drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments are some, but not all, embodiments of this application. The embodiments described below with reference to the accompanying drawings are exemplary and intended to explain this application, and should not be construed as limiting this application. All other embodiments obtained by those skilled in the art based on the embodiments of this application without creative effort are within the scope of protection of this application. The embodiments of this application will be described in detail below with reference to the accompanying drawings.

[0079] In the description of this application, it should be understood that the terms "center", "longitudinal", "lateral", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limiting the scope of protection of this application.

[0080] The following is in conjunction with the appendix Figures 1 to 8 This application will be described in further detail.

[0081] This application provides a method for designing wing structural parameters that considers the load-bearing efficiency of the wing cover, such as... Figure 1 As shown, it includes the following steps:

[0082] Step 1: Obtain the initial load-bearing coefficient of the hatches at different locations on the wing;

[0083] Step 2: Construct a finite element model of the wing structure and correct the Young's modulus of the wing cover material in the finite element model based on the initial load-bearing coefficient;

[0084] Step 3: Optimize the wing structure parameters based on the finite element model of the wing structure to obtain the wing structure parameters. Extract the stress level during the optimization process and obtain the new load-bearing coefficient of the hatch at this stress level. Compare the new load-bearing coefficient with the initial load-bearing coefficient. If the convergence condition is not met, return to Step 2. Correct the Young's modulus of the hatch material in the finite element model of the wing structure based on the new load-bearing coefficient until the wing structure parameters that meet the convergence condition are obtained.

[0085] The wing structural parameter design method considering the load-bearing efficiency of the wing cover in this application, in step one, obtains the initial load-bearing coefficient of the wing cover at different locations, including:

[0086] S101. Determine the method for obtaining the bearing capacity coefficient of the cover under different stress levels;

[0087] S102. Obtain the stress level of the cover area at different locations on the wing;

[0088] S103. Determine the initial load-bearing coefficient of the cover at different positions of the wing based on the stress level.

[0089] Based on the overall wing structure layout, typical structures are selected for testing or simulation to obtain the differences in tensile and compressive load transmission between the hatch and the non-hatch area of ​​the wall panel, thereby obtaining the hatch's load-bearing coefficient. Since the actual load transmission of the hatch is related to fastener clearance, fastener preload, and friction in the hatch contact area, a relatively complex detailed model needs to be established for analysis. Simplifications in numerical simulations may affect the results. To obtain a more accurate hatch load transmission efficiency, experimental methods are recommended. Tests can be conducted at the box segment level or stiffened plate level, with test specimen forms such as... Figure 2 As shown, all tests were conducted with one end loaded and the other end fixed. The difference was that the stiffened plate level test required setting up dummy parts at the ribs to simulate rib support.

[0090] In one embodiment of this application, in S101, the load-bearing coefficient of the cap under different stress levels is obtained by using a box segment-level test, including:

[0091] Obtain box-level test specimens for box-level testing;

[0092] Box segment level tests were conducted under different stress levels to obtain box segment level test data;

[0093] The load-bearing coefficient is calculated based on the test data of the box section.

[0094] like Figure 3 As shown, the box-section test specimen includes: a wing beam 107, a first cover 103, a first cover strain gauge 104, and a first skin strain gauge 105. The wing beam 107 has an opening in its wall panel 102, and a fixed end 106 is provided at the end of the wing beam 107. The first cover 103 is installed at the opening in the wall panel. The first cover strain gauge 104 is disposed on the first cover 103. The first skin strain gauge 105 is disposed on the wall panel 102, and both the first skin strain gauge 105 and the first cover strain gauge 104 are located on the symmetry axis 101 of the cover.

[0095] In another embodiment of this application, in S101, the bearing capacity coefficient of the cap under different stress levels is obtained by using a stiffened plate level test, including:

[0096] Obtain stiffened plate-level test specimens for stiffened plate-level testing;

[0097] Stiffened plate level tests were conducted under different stress levels to obtain stiffened plate level test data;

[0098] The bearing capacity coefficient is calculated based on the test data of the stiffened slab.

[0099] like Figure 4 As shown, the stiffened plate-level test specimen includes: a stiffened plate 108, a second cover 110, a second cover strain gauge 111, and a second skin strain gauge 112. The stiffened plate 108 has a stiffened plate opening, and dummy support pieces are provided at the ribs 113 of the stiffened plate 108 to simulate rib support. A fixed end 109 is provided at the end of the stiffened plate 108. The second cover 110 is installed at the stiffened plate opening. The second cover strain gauge 111 is disposed on the second cover 110. The second skin strain gauge 112 is disposed on the stiffened plate 108, and both the second skin strain gauge 112 and the second cover strain gauge 111 are parallel to the ribs 113.

[0100] The bearing capacity factor is calculated based on test data at the box section level or the stiffened slab level, including:

[0101] The calculated stress level is:

[0102] ;

[0103] in, For skin stress level, E sk The Young's modulus of the skin material. The strain value of the skin strain gauge. This refers to the number of skin strain gauges;

[0104] Calculate the load-bearing capacity of the cover at this stress level, including:

[0105] ;

[0106] in, For bearing capacity, The strain value of the strain gauge on the lid. The number of strain gauges for the lid. For the thickness of the cap, This refers to the skin thickness.

[0107] Based on the bending moment M of the airfoil section at the hatch, and considering the local axial compressive stability of the skin and the allowable compressive stress constraints of the panel material, the stress level at the hatch is determined. The specific calculation of the skin thickness under the control of local axial compressive stability is as follows:

[0108] ;

[0109] Meanwhile, taking into account the minimum processing thickness t of the skin min The specific calculation of the skin stress level under the control of local axial compression stability is as follows:

[0110] In this embodiment, S102, obtaining the stress level of the cover area at different positions of the wing includes:

[0111] Calculate the skin stress level:

[0112] ;

[0113] in, For skin stress level, E is the critical compressive stress coefficient. sk Let M be the Young's modulus of the skin material, M be the bending moment of the airfoil section at the hatch, μ be the Poisson's ratio of the skin material, b be the width of the loaded edge of the skin, and A be the bending moment of the airfoil section at the hatch. box Let t be the cross-sectional area of ​​the wing box section at the hatch, λ be the ratio of the area of ​​the stringer to the area of ​​the skin section, and t be the cross-sectional area of ​​the wing box section at the hatch. min This refers to the minimum processing thickness of the skin.

[0114] Considering the allowable compressive stress of the wall panel material, the stress level in the cover area is determined based on the skin stress level:

[0115] ;

[0116] in, The stress level in the lid area, This is the stress level reduction factor. This represents the allowable stress of the wall panel.

[0117] The allowable stress value of the panel is usually determined by factors such as aircraft materials and service life, and can be finally determined based on relevant material test data. The stress level reduction factor may be overestimated because the estimation of the skin stress level only considers the local axial compression stability constraint of the skin and does not take into account factors such as the strength of the panel columns and the composite stability of the skin compression and shear. Different stress level reduction factor values ​​can be used for different areas of the cover. It is recommended that the stress level reduction factor value be larger as it is closer to the wing root.

[0118] In S103, the stress level is calculated, and the initial load-bearing coefficient of the wing opening at different positions is determined according to the test method in S101.

[0119] The wing structure parameter design method considering the load-bearing efficiency of the wing cover in this application, in step two, corrects the Young's modulus of the wing cover material in the finite element model of the wing structure based on the initial load-bearing coefficient, including:

[0120] For the hatch at the i-th wing position, the corrected Young's modulus of the hatch material is:

[0121] ;

[0122] Among them, E i Let E be the Young's modulus of the cover material after correction at the i-th wing position. Let be the initial load-bearing coefficient of the cover at the i-th wing position.

[0123] like Figure 5 As shown, the Young's modulus of the cover material is adjusted in the finite element model of the wing structure to approximate the actual load-bearing capacity of the cover, and the load-bearing coefficient of the cover is substituted into the finite element model of the wing structure.

[0124] The wing structural parameter design method considering the load-bearing efficiency of the hatch in this application, in step three, obtains the basic structural parameters of the final skin and hatch area through wing structural parameter optimization design. During the optimization process, a new stress level in the hatch area is extracted simultaneously, and a new load-bearing coefficient for the hatch area is obtained using the method in step one based on this new stress level. Then return to step two to update the material of the cap area unit and continue optimization until the desired result is achieved. Requirements such as strength, rigidity, stability, and manufacturability are considered. The final thickness of the skin on both sides of the cap is obtained. , Cap thickness Area of ​​the long trusses on both sides of the lid , ,like Figure 6 As shown.

[0125] The wing structure parameter design method considering the load-bearing efficiency of the cover in this application also includes step five: reinforcing and optimizing the structure of the cover area to obtain the reinforced and optimized wing structure parameters.

[0126] like Figure 7 As shown, the optimized wing structure parameters are as follows:

[0127] ;

[0128] ;

[0129] ;

[0130] ;

[0131] in, , To reinforce the thickness of the skin on both sides of the cap after optimization, , The thickness of the skin on both sides of the flap, b is the skin reinforcement coefficient. p b1 and b2 are the widths of the opening and the skin reinforcement, respectively. , To compensate for the area of ​​the stringers on both sides of the optimized cover, , The area of ​​the long stringers on both sides of the lid, This is the reinforcement coefficient for the stringer.

[0132] The skin and stringer reinforcement factor is related to the opening type, aircraft load level, and service life requirements. Its specific value needs to be defined based on the actual aircraft conditions. The method for determining the value is as follows: ,and .

[0133] In one specific embodiment of this application, a test specimen for the load-bearing efficiency of the box-section panel cover is designed according to the wing structure arrangement, and the load-bearing coefficient of the cover under different stress levels is obtained through experiments. Figure 2 As shown, to obtain a more realistic stress level in the opening area, at least three stringers should be retained on one side of the chordal opening, depending on the location of the opening. If the width of the wing box prevents the retention of at least three stringers, the actual structural design can be followed. A preliminary assessment of the stress level in the wing panel opening area is conducted, and the load-bearing coefficient for this area is obtained. Based on the load-bearing coefficient, the Young's modulus of the material in the opening area of ​​the finite element model is reduced. The model is shown below. Figure 5 As shown; follow the procedure of this application to complete parameter optimization and obtain the basic structural parameters that meet the design requirements (including load-bearing efficiency convergence, strength, stiffness, stability, and manufacturing process, etc.), as shown. Figure 6 As shown in the diagram. Finally, the design was reinforced and optimized; a schematic diagram of the thickened area of ​​the wall panel is shown below. Figure 7 As shown.

[0134] In this embodiment, the skin reinforcement coefficient Take 0.5 as the stringer reinforcement coefficient. Taking 0.35, we finally get , as follows, and The following formula can be used to calculate the final numerical model of the opening region structure. Figure 8 As shown.

[0135] ;

[0136] .

[0137] The wing structure parameter design method considering the load-bearing efficiency of the cover in this application can obtain the actual load-bearing capacity of the cover and quickly obtain the actual load-bearing capacity of the cover under different stress levels; it corrects the Young's modulus of the cover material in the finite element model, making the load transfer of the panel in the finite element model more consistent with the actual situation; it can optimize the basic structural parameters of the wing panel, and simultaneously consider strength, process and actual load-bearing capacity of the cover while optimizing and iterating the basic structural parameters of the wing panel, resulting in more reliable structural parameters and avoiding a lot of corrections later; through the detailed parameter design of the cover area of ​​the panel, it also takes into account stress concentration in the opening area and stiffness enhancement in the frame area, thereby increasing the strength and life of the cover area while increasing sealing reliability.

[0138] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. A method for designing wing structural parameters considering the load-bearing efficiency of the flap, characterized in that, include: Step 1: Obtain the initial load-bearing coefficient of the hatches at different locations on the wing; Step 2: Construct a finite element model of the wing structure and correct the Young's modulus of the wing cover material in the finite element model based on the initial load-bearing coefficient; Step 3: Optimize the wing structure parameters based on the finite element model of the wing structure to obtain the wing structure parameters. Extract the stress level during the optimization process and obtain the new load-bearing coefficient of the hatch at this stress level. Compare the new load-bearing coefficient with the initial load-bearing coefficient. If the convergence condition is not met, return to Step 2. Correct the Young's modulus of the hatch material in the finite element model of the wing structure based on the new load-bearing coefficient until the wing structure parameters that meet the convergence condition are obtained.

2. The wing structure parameter design method considering the load-bearing efficiency of the hatch according to claim 1, characterized in that, In step one, the initial load-bearing coefficients of the openings at different locations on the wing are obtained, including: S101. Determine the method for obtaining the bearing capacity coefficient of the cover under different stress levels; S102. Obtain the stress level of the cover area at different locations on the wing; S103. Determine the initial load-bearing coefficient of the wing openings at different positions based on the stress level.

3. The wing structure parameter design method considering the load-bearing efficiency of the canopy according to claim 2, characterized in that, In S101, the load-bearing coefficient of the cover under different stress levels is obtained using box segment-level tests, including: Obtain box-level test specimens for box-level testing; Box segment level tests were conducted under different stress levels to obtain box segment level test data; The load-bearing coefficient is calculated based on the test data of the box section.

4. The wing structure parameter design method considering the load-bearing efficiency of the hatch according to claim 3, characterized in that, Box segment-level test specimens include: A wing beam (107) has a wall panel (102) with an opening in the wall panel and a fixed end (106) at the end of the wing beam (107). The first cover (103) is installed at the opening of the wall panel; The first strain gauge (104) is disposed on the first cover (103); The first skin strain gauge (105) is disposed on the wall panel (102), and the first skin strain gauge (105) and the first cover strain gauge (104) are both located on the cover symmetry axis (101).

5. The wing structure parameter design method considering the load-bearing efficiency of the hatch according to claim 2, characterized in that, In S101, the load-bearing coefficient of the cap under different stress levels is obtained using stiffened plate level tests, including: Obtain stiffened plate-level test specimens for stiffened plate-level testing; Stiffened plate level tests were conducted under different stress levels to obtain stiffened plate level test data; The bearing capacity coefficient is calculated based on the test data of the stiffened slab.

6. The wing structure parameter design method considering the load-bearing efficiency of the canopy according to claim 5, characterized in that, The stiffened plate-level test specimens include: A stiffening plate (108) is provided with a stiffening plate opening, a dummy wing rib support is provided at the wing rib (113) of the stiffening plate (108), and a fixed end (109) is provided at the end of the stiffening plate (108). The second cover (110) is installed at the opening of the stiffening plate; The second strain gauge (111) is disposed on the second cover (110); The second skin strain gauge (112) is disposed on the stiffening plate (108), and both the second skin strain gauge (112) and the second cover strain gauge (111) are parallel to the wing rib (113).

7. The wing structure parameter design method considering the load-bearing efficiency of the hatch according to claim 3 or 6, characterized in that, The calculated stress level is: ； in, For skin stress level, E sk The Young's modulus of the skin material. The strain value of the skin strain gauge. This refers to the number of skin strain gauges; Calculate the load-bearing capacity of the cover at this stress level, including: ； in, For bearing capacity, The strain value of the strain gauge on the lid. The number of strain gauges for the lid. For the thickness of the cap, This refers to the skin thickness.

8. The wing structure parameter design method considering the load-bearing efficiency of the hatch according to claim 7, characterized in that, In S102, the stress levels of the wing canopy area at different locations are obtained, including: Calculate skin thickness: ； Calculate the skin stress level: ； in, For skin stress level, E is the critical compressive stress coefficient. sk Let M be the Young's modulus of the skin material, M be the bending moment of the airfoil section at the hatch, μ be the Poisson's ratio of the skin material, b be the width of the loaded edge of the skin, and A be the bending moment of the airfoil section at the hatch. box Let t be the cross-sectional area of ​​the wing box section at the hatch, λ be the ratio of the area of ​​the stringer to the area of ​​the skin section, and t be the cross-sectional area of ​​the wing box section at the hatch. min This refers to the minimum processing thickness of the skin. Determine the stress level in the flap area based on the skin stress level: ； in, The stress level in the lid area, This is the stress level reduction factor. This represents the allowable stress of the wall panel.

9. The wing structure parameter design method considering the load-bearing efficiency of the hatch according to claim 8, characterized in that, In step two, the Young's modulus of the wing cover material in the finite element model of the wing structure is corrected based on the initial load-bearing coefficient, including: For the hatch at the i-th wing position, the corrected Young's modulus of the hatch material is: ； Among them, E i Let E be the Young's modulus of the cover material after correction at the i-th wing position. Let be the initial load-bearing coefficient of the cover at the i-th wing position.

10. The wing structure parameter design method considering the load-bearing efficiency of the hatch according to claim 9, characterized in that, In step three, the convergence conditions include: ； in, Let be the new load-bearing coefficient for the cover at the i-th wing position.

11. The wing structure parameter design method considering the load-bearing efficiency of the canopy according to claim 10, characterized in that, In step three, the convergence conditions also include: strength, stiffness, stability, and manufacturability requirements.

12. The wing structure parameter design method considering the load-bearing efficiency of the hatch according to claim 11, characterized in that, In step three, the wing structure parameters include: the thickness of the skin on both sides of the hatch, the thickness of the hatch, and the area of ​​the stringers on both sides of the hatch.

13. The wing structure parameter design method considering the load-bearing efficiency of the hatch according to claim 12, characterized in that, It also includes step five, which involves reinforcing and optimizing the structure of the wing cover area to obtain the reinforced and optimized wing structure parameters.

14. The wing structure parameter design method considering the load-bearing efficiency of the hatch according to claim 13, characterized in that, The optimized and reinforced wing structure parameters are as follows: ； ； ； ； in, , To reinforce the thickness of the skin on both sides of the cap after optimization, , The thickness of the skin on both sides of the flap, b is the skin reinforcement coefficient. p b1 and b2 are the widths of the opening and the skin reinforcement, respectively. , To compensate for the area of ​​the stringers on both sides of the optimized cover, , The area of ​​the long stringers on both sides of the lid, This is the reinforcement coefficient for the stringer.

15. The wing structure parameter design method considering the load-bearing efficiency of the canopy according to claim 14, characterized in that, The reinforcement coefficient is determined as follows: ,and .

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