A method for determining the equivalent stiffness of the boundary of an aircraft panel structure

Abstract

The present invention discloses a method for determining the equivalent boundary stiffness of an aircraft panel structure, belonging to the technical field of aircraft structural strength. This method respectively conducts finite element modeling on the fuselage and the local panel, applies loads at the loading center points of the two finite element models, obtains the corresponding displacement-load curves and boundary rotation-load curves. After comparing the two displacement-load curves and the two boundary rotation-load curves, the total weighted error is obtained. By updating the boundary stiffness of the local panel in the finite element model, the total weighted error meets the requirements, and the finally obtained boundary stiffness is used as the equivalent boundary stiffness. This method realizes the consistency of the deformation mode and failure behavior between the local panel structure and the overall fuselage. The small-size local panel structure can be used to replace the large-size fuselage frame section structure, which can significantly improve the analysis efficiency and reduce the test cost at the same time.

Description

Technical Field

[0001] The present invention belongs to the technical field of aircraft structural strength, and particularly relates to a method for determining the equivalent stiffness of the boundary of an aircraft panel structure. Background Art

[0002] In the field of civil aviation, ground service equipment within airports plays a crucial role, including but not limited to various airport ground support vehicles, boarding bridges, gangways, maintenance gantries, and baggage handling devices. These devices play an indispensable role in the docking, maintenance of aircraft, as well as the boarding of passengers and the loading of baggage. However, during the docking stage of an aircraft, due to reasons such as improper operation, equipment failure, or environmental factors, such ground service equipment may collide with the aircraft. This accidental collision between ground service equipment and the fuselage structure has become one of the main factors leading to structural damage of civil aircraft, posing a serious threat to flight safety.

[0003] With the continuous progress of aviation technology, the design of new-generation aircraft adopts a large number of composite panel structures, which perform excellently in reducing the weight of the aircraft, improving fuel efficiency, and enhancing the overall performance of the aircraft. However, when such composite panel structures are subjected to impact, their performance is quite different from that of traditional metal materials. Under the action of impact, the composite panel may not show obvious structural deformation, dents, or cracks and other visually visible damages, but the impact may cause serious internal damages to the structure. These internal damages include but are not limited to the damage of the fuselage frame, the debonding of stringers or shear bands, the shear failure of bolts, and the fracture of stringers. These damage forms not only seriously reduce the structural performance of the aircraft but may also pose a direct threat to flight safety.

[0004] In view of the serious consequences that may be brought about by ground service equipment impacting the aircraft, in order to ensure the service safety of the aircraft, it is necessary to conduct in-depth research on the relevant situations of ground service equipment impacting the aircraft. However, during the actual research process, due to considerations such as cost, safety, and feasibility, it is obviously unrealistic to directly use the aircraft fuselage for direct impact tests. Therefore, many studies have begun to focus on the simulation process, attempting to reproduce the impact process of the aircraft through computer simulation to evaluate the impact on the aircraft structure.

[0005] Although simulation technology is increasingly widely used in the aviation field, it still faces many challenges when simulating the specific scenario of ground service equipment impacting the aircraft. Due to the complexity and uncertainty of the impact process, it is difficult for simulation to reflect the impact on the aircraft fuselage or aircraft panel during the actual impact process. This results in the difficulty of ensuring the accuracy of the simulation results, thus limiting the application of simulation technology in evaluating aircraft structural damage.

[0006] On the other hand, if the whole machine is directly used for the impact test, although more accurate data can be obtained, it is obviously costly and difficult to implement. This is not only because the whole machine test requires a large amount of resources and funds, but also because of the potential safety risks and environmental impacts during the test process. Summary of the Invention

[0007] The purpose of the present invention is to overcome the above-mentioned disadvantages of the prior art and provide a method for determining the equivalent stiffness of the boundary of an aircraft panel structure, so as to solve the problem that the existing technology has inaccurate test simulation methods for the impact of ground service equipment on the aircraft and it is difficult to evaluate the influence of low-speed impact on the aircraft structure damage.

[0008] To achieve the above purpose, the present invention adopts the following technical solutions:

[0009] A method for determining the equivalent stiffness of the boundary of an aircraft panel structure includes the following steps:

[0010] S1. Perform finite element modeling on the fuselage to obtain a fuselage finite element model; apply a load at the first loading center point of the fuselage, gradually increase the load until the set load is reached, and obtain the first displacement-load curve and the first boundary rotation-load curve of the first loading center point;

[0011] S2. Perform finite element modeling on the local panel to obtain a local panel finite element model; the local panel is a partial arc plate of the fuselage; simplify the structure of the local panel, calculate and obtain the boundary stiffness of the local panel as the initial boundary condition of the local panel finite element model, apply a load at the second loading center point of the local panel, gradually increase the load until the set load is reached, and obtain the second displacement-load curve and the second boundary rotation-load curve of the second loading center point;

[0012] S3. Compare the first displacement-load curve and the second displacement-load curve to obtain the first weighted error; compare the first boundary rotation-load curve and the second boundary rotation-load curve to obtain the second weighted error, and combine the first weighted error and the second weighted error to obtain the total weighted error;

[0013] S4. Starting from the initial boundary condition, update the boundary stiffness through an optimization method, and calculate the second displacement-load curve and the second boundary rotation-load curve of the second loading center point based on the updated boundary stiffness;

[0014] S5. Repeat S3 and S4 until the total weighted error is reduced to the set range, and the finally obtained boundary stiffness is the equivalent boundary stiffness;

[0015] S6. Based on the equivalent boundary stiffness, conduct an impact test of the ground service equipment on the aircraft panel structure for the local panel.

[0016] A further improvement of the present invention is:

[0017] Preferably, in S1, the fuselage is an integral fuselage or a 1 / 4 fuselage structure.

[0018] Preferably, in S1, when the fuselage is a 1 / 4 fuselage structure, the boundary condition of the fuselage finite element model is a symmetric boundary condition.

[0019] Preferably, in S2, the local panel is a panel structure including three frames and five stringers, or a panel structure including three frames and six stringers.

[0020] Preferably, in S1, the finite element modeling of the fuselage includes the following steps:

[0021] S101, set the constitutive parameters and damage parameters of the fuselage material, set the ply direction, and obtain the fuselage model;

[0022] S102, perform geometric simplification and shelling on the fuselage model;

[0023] S103, divide the mesh of the fuselage model after geometric simplification and shelling to obtain the meshed fuselage model;

[0024] S104, establish the skin, stringers, frames and shear bands in the meshed fuselage model, as well as their connection relationships;

[0025] S105, apply boundary conditions to the fuselage model to obtain the fuselage finite element model.

[0026] Preferably, in S2, the process of simplifying the structure of the local panel is as follows: simplify the frames in the local panel into arcs, simplify the skin into multiple arcs, the arc directions of the simplified skin are the same as those of the simplified frames, the arc radian of the simplified skin is the same as that of the simplified frames, and the stringers and shear bands in the local panel are ignored.

[0027] Preferably, in S2, the boundary stiffness includes rotational stiffness, longitudinal stiffness and radial stiffness.

[0028] Preferably, in S1 and S2, the initial value of the applied load and the step size of the increased load are equal.

[0029] Preferably, in S3, the calculation formula for the total weighted error is:

[0030]

[0031] Where is the total weighted error, and are the displacements of the loading points calculated from the fuselage finite element model and the local panel finite element model respectively, and are the boundary rotation angles calculated from the fuselage finite element model and the local panel finite element model respectively, and are the corresponding weighting coefficients.

[0032] Preferably, in S5, the setting range ≤ 0.01.

[0033] Compared with the prior art, the present invention has the following beneficial effects:

[0034] The present invention discloses a method for determining the boundary equivalent stiffness of an aircraft panel structure. The method respectively performs finite element modeling on the fuselage and the local panel, applies loads at the loading center points in the two finite element models, respectively obtains the corresponding displacement-load curves and boundary rotation angle-load curves, obtains the first weighted error of the two displacement-load curves and the second weighted error of the two boundary rotation angle-load curves, obtains the total weighted error based on the first weighted error and the second weighted error, and further updates the boundary stiffness of the local panel in the finite element model so that the total weighted error meets the requirements. The finally obtained boundary stiffness is determined as the boundary equivalent stiffness. This method combines the simplified theoretical method and the refined finite element method, can accurately and quickly calculate the true boundary stiffness characteristics of the aircraft local panel through the finite element model, apply it to the local panel, and ensure that the boundary stiffness characteristics of the local panel are consistent with those when impacting the whole fuselage, realizing the consistency of the deformation mode and failure behavior between the local panel structure and the whole fuselage. Thus, the local panel structure can be used to carry out impact analysis and tests, while reducing the analysis and test costs, ensuring the authenticity and reliability of the anti-impact characteristic analysis and test results of the fuselage composite panel structure. BRIEF DESCRIPTION OF THE DRAWINGS

[0035] Figure 1 is a flow chart of the method for determining the boundary equivalent stiffness of an aircraft panel structure according to the present invention;

[0036] Figure 2 is a schematic diagram of the positional relationship between the whole fuselage and the local panel structure of the present invention;

[0037] Figure 3 is a schematic diagram of the positional relationship between the 1 / 4 fuselage structure and the local panel structure of the present invention;

[0038] Figure 4 is the front view of the local panel structure;

[0039] Figure 5 is the three-dimensional side view of the local panel structure;

[0040] Figure 6 is the finite element model diagram of the 1 / 4 fuselage structure;

[0041] Figure 7 is the schematic diagram of the boundary stiffness simulation;

[0042] Among them, 1. the overall fuselage; 2. the 1 / 4 fuselage structure; 3. the local panel; 4. the frame; 5. the stringer; 6. the skin. Specific implementation manner

[0043] Hereinafter, the terms "first", "second", "third", and "fourth" are only used for descriptive purposes and cannot be understood as indicating or implying relative importance or implicitly specifying the quantity of the indicated technical features. Thus, the features defined with "first", "second", "third", and "fourth" may explicitly or implicitly include one or more of such features.

[0044] It should be noted that the terms "first", "second", etc. in the description and drawings of the present invention are used to distinguish similar objects and do not necessarily need to describe a specific order or sequence. It should be understood that such data can be interchanged under appropriate circumstances so that the embodiments of the present invention described herein can be implemented in an order other than those illustrated or described herein. In addition, the terms "comprising" and "having" and any variations thereof are intended to cover non-exclusive inclusion. For example, a process, method, system, product, or device comprising a series of steps or units does not necessarily have to be limited to those steps or units clearly listed, but may include other steps or units not clearly listed or inherent to these processes, methods, products, or devices.

[0045] The method of the present invention discloses a method for determining the equivalent stiffness of the boundary of an aircraft panel structure, including the following steps:

[0046] S1. Perform finite element modeling on the fuselage to obtain a fuselage finite element model; apply a load at the first loading center point of the fuselage, gradually increase the load until the set load is reached, and obtain the first displacement-load curve and the first boundary rotation-load curve of the first loading center point;

[0047] S2. Perform finite element modeling on the local panel 3 to obtain a local panel 3 finite element model; the local panel 3 is a partial arc plate on the fuselage; simplify the structure of the local panel 3, calculate and obtain the boundary stiffness of the local panel 3 as the initial boundary condition of the local panel 3 finite element model, apply a load at the second loading center point of the local panel 3, gradually increase the load until the set load is reached, and obtain the second displacement-load curve and the second boundary rotation-load curve of the second loading center point;

[0048] S3. Compare the first displacement-load curve and the second displacement-load curve to obtain the first weighted error; compare the first boundary rotation-load curve and the second boundary rotation-load curve to obtain the second weighted error, and combine the first weighted error and the second weighted error to obtain the total weighted error;

[0049] S4. Starting from the initial boundary conditions, update the boundary stiffness through an optimization method, and calculate the second displacement load curve and the second boundary rotation angle load curve of the second loading center point based on the updated boundary stiffness;

[0050] S5. Repeat S3 and S4 until the total weighted error is reduced to the set range, and the finally obtained boundary stiffness is the boundary equivalent stiffness;

[0051] S6. Based on the boundary equivalent stiffness, conduct a ground service equipment impact test on the aircraft panel structure for the local panel 3.

[0052] The present invention uses the local panel 3 to replace the structure of the integral fuselage 1, and proposes an effective method and a feasible process for calculating the boundary stiffness of the local panel 3, so as to achieve the same deformation mode and failure behavior between the local panel 3 and the integral fuselage 1, ensure the reliability of the analysis or test results, and at the same time can significantly reduce the cost and difficulty of the analysis and test; this method combines a simplified theoretical method and a refined finite element method, quickly calculates an estimated value of the boundary stiffness of the local panel 3 through a theoretical algorithm, and uses it as the initial value of the boundary stiffness of the finite element model of the local panel 3 to perform numerical iterative calculations to obtain the optimized boundary stiffness value, which not only improves the calculation efficiency but also ensures the accuracy of the stiffness calculation.

[0053] In some embodiments of the present invention, the fuselage in S1 is an integral fuselage 1 or a 1 / 4 fuselage structure 2. Refer to Figure 2 for the structural relationship between the integral fuselage 1 and the local panel 3. This figure is a schematic diagram. The actual length of the integral fuselage 1 is much longer than the length in this schematic diagram. Preferably, the local panel 3 is arranged at the central position in the length direction of the integral fuselage 1.

[0054] Refer to Figure 3 , the 1 / 4 fuselage structure 2 is a curved surface, and the curved surface is composed of a continuous array of arcs; the arc length of the arc is 1 / 4 of the circumference, and this arc is an arc segment on the cross-section perpendicular to the fuselage axis, and the corresponding radian of this arc is 90°. Preferably, the local panel 3 in the 1 / 4 fuselage structure 2 is still at the central position in the length direction of the 1 / 4 fuselage structure 2.

[0055] Through Figure 2 and Figure 3 it can be seen that the local panel 3 is a part of the integral fuselage 1 or the 1 / 4 fuselage structure 2 and is an arc-shaped plate. By restricting the edge of the local panel 3 during impact through the equivalent stiffness, it is equivalent to that during the impact test, the local panel 3 is still in the integral fuselage 1 or the 1 / 4 fuselage structure 2, making the impact test results true and reliable.

[0056] Preferably, finite element modeling is performed on the 1 / 4 fuselage structure 2 to replace the integral fuselage 1, which can simplify the simulation process and accelerate the solution efficiency.

[0057] In some embodiments of the present invention, referring to Figure 4 and Figure 5 , in the overall fuselage 1 structure of the aircraft, considering the local effects according to the selected impact area, the structure and size of the local panel 3 are selected. An arc-shaped panel including three frames 4 and five stringers 5 is selected as the local panel 3, or an arc-shaped panel including three frames 4 and six stringers 5 is selected as the local panel 3; while ensuring the impact test effect, the model can be simplified.

[0058] Referring to Figure 4 and Figure 5 , Figure 4 is the front view of the local panel 3, Figure 5 is the three-dimensional side view of the local panel 3. The local panel 3 includes a skin 6, stringers 5 and frames 4. The stringers 5 are arranged along the axis direction of the whole fuselage, and the frames 4 are arranged along the circumferential direction of the whole fuselage. The stringers 5 and the frames 4 are connected by shear bands, and the planes where the stringers 5 and the frames 4 are located are perpendicular to each other; it should be understood that the complete frame 4 is a circumferential line, and only a part of it is shown in the figure.

[0059] In some embodiments of the present invention, in S1, the finite element modeling of the fuselage includes the following steps:

[0060] S101, set the constitutive parameters and damage parameters of the fuselage, set the ply thickness and ply direction to obtain the fuselage model;

[0061] S102, perform geometric simplification and shelling on the fuselage model;

[0062] S103, perform mesh division on the fuselage model after geometric simplification and shelling. The mesh is mainly composed of quadrilateral meshes and supplemented by triangular meshes to obtain the meshed fuselage model;

[0063] S104, establish the skin 6, stringers 5, frames 4 and shear bands, and their connection relationships in the meshed fuselage model;

[0064] S105, apply boundary stiffness to the fuselage model. It should be noted that the boundary conditions applied to the overall fuselage 1 are to apply boundary conditions to both sides. If the fuselage is a 1 / 4 fuselage structure 2, the applied boundary conditions are two symmetric boundary conditions on the top, bottom, left and right; after applying the boundary conditions, the finite element modeling of the whole fuselage is completed.

[0065] In some embodiments of the present invention, in S2, the finite element modeling process of the local panel 3 is the same as the above-mentioned finite element modeling process of the fuselage, but the applied boundary conditions are different.

[0066] Specifically, referring to Figure 6During the finite element modeling process of the local panel 3, according to the structure of the local panel 3, and based on its size and structural composition, a simplified theoretical method is adopted for simplification. The skin 6 and the frame 4 can be simplified into one or more arcs with a rectangular cross-section; the skin 6 is simplified into multiple arcs with a rectangular cross-section and arranged in an array, and the frame 4 itself is simplified into an arc with a rectangular cross-section and a certain thickness; the arc direction of the simplified skin 6 is the same as that of the simplified frame 4, and the arc curvature of the simplified skin 6 is the same as that of the simplified frame 4. Since the stringer 5 and the shear band have little influence on the boundary stiffness, they can be ignored during the simulation calculation.

[0067] Preferably, referring to Figure 7 , the boundary condition applied to the local panel 3 is the boundary stiffness, and the initial boundary condition applied is the calculated boundary stiffness, which is an inherent attribute value of the local panel 3. The boundary stiffness includes rotational stiffness , longitudinal stiffness and radial stiffness ; the rotational stiffness is the resistance bending stiffness at the boundary of the local panel 3 when subjected to a horizontal radial load along the arc (this stiffness is represented by a spiral line in the figure); the longitudinal stiffness is the resistance translation stiffness in the direction perpendicular to the load at the boundary when subjected to a horizontal radial load along the arc; the radial stiffness is the resistance translation stiffness at the boundary parallel to and opposite to the load direction when subjected to a horizontal radial load along the arc. According to the actual situation during impact, the rotational stiffness is decomposed into the stiffness when the upper and lower edges are stressed, the radial stiffness is decomposed into the stiffness when the upper and lower edges are stressed, and the longitudinal stiffness mainly relies on the bottom support. Therefore, only the constraint of the radial stiffness is set at the bottom. In the figure, F is the abbreviation of the force applied by the horizontal radial load.

[0068] Through the theoretical calculation formulas of the three boundary stiffnesses, the values of the three boundary stiffnesses are calculated as the initial boundary conditions for subsequent iterative calculations. The calculation formulas of the rotational stiffness , longitudinal stiffness and radial stiffness are shown in the following formulas (1), (2) and (3) respectively:

[0069]

[0070]

[0071]

[0072] Among them, is the elastic modulus, is the Poisson's ratio, is the width of the arc cross-section, is the equivalent thickness of the arc cross-section, is the arc length, is the arc radius. When performing the simulation calculation on the local panel 3, relevant parameters of the local panel 3 are substituted to calculate the boundary stiffness conditions of the local panel 3 model 、 and initial values.

[0073] It should be noted that the above formula is only applicable to the assumption that the cross-sectional dimensions of the arc along the axial direction are uniform. Some components in the actual model are not within its scope of use. Therefore, the calculated boundary stiffness value will deviate from the actual value to a certain extent and cannot be used as the final result, but can only be used as the input value for the subsequent automated parameter optimization process.

[0074] In some embodiments of the present invention, loading areas are provided in both the fuselage in S1 and the local panel 3 in S2. The loading areas are all actual impact areas and are part of the fuselage model or the local panel 3. In actual tests, the impact area refers to the area where the anti-collision strip of airport ground service equipment contacts the skin 6, generally a slender rectangular area, whose length generally does not exceed 2000 mm and width generally does not exceed 100 mm. The first loading center point in S1 is located in the loading area of the fuselage, and the second loading center point in S2 is located in the loading area of the local panel 3. As a preferred solution, both loading center points are at the geometric position center of the actual loading area.

[0075] In S1, the first loading center point is the actual impact point on the fuselage that is vulnerable to impact or the simulated position. The first displacement load curve therein represents the curve of the load change and the position change of the first loading center point, and the first boundary rotation angle load curve represents the relationship between the load change at the first loading center point and the boundary rotation angle of the loading area, and this boundary is the boundary of the loading area.

[0076] In some embodiments of the present invention, in S2, the second loading center point is the actual impact point of the local panel 3 during the test. The second displacement load curve therein represents the curve of the load change and the position change of the second loading center point, and the second boundary rotation angle load curve represents the relationship between the load change at the second loading center point and the boundary rotation angle of the loading area, and this boundary is the boundary of the loading area.

[0077] It should be noted that in S1 and S2, the initial values of the loading loads and the step sizes of the increased loads are the same, so that during the simulation process, the local panel 3 and the fuselage are simulated under the same working condition loads.

[0078] In some embodiments of the present invention, in S3, after the load is loaded to the set value, the displacement of the loading point of the fuselage and the local panel 3 and the boundary rotation angle are compared, and two weighted errors, namely the first weighted error and the second weighted error, are calculated. The first weighted error and the second weighted error are respectively multiplied by the corresponding coefficients and then added to obtain the total weighted error. At the beginning of the calculation, it is difficult for the total weighted error to meet the set range value, indicating that the loading stiffness of the local panel 3 is quite different from the simulation of the overall fuselage 1 or the 1 / 4 fuselage structure 2. Therefore, the boundary stiffness value is updated based on the generalized gradient descent method for iterative calculation.

[0079] Among them, the calculation formula for the weighted error is:

[0080] (4)

[0081] Wherein, and are respectively the displacements of the loading points calculated by the finite element model of the fuselage and the finite element model of the local panel 3, and are respectively the boundary rotation angles calculated by the finite element model of the fuselage and the finite element model of the local panel 3, and are the corresponding weighted coefficients, and The value ranges of both are 0 - 1, and + = 1, and the default value range is = = 0.5, and specific adjustments can be made according to the requirements for rotational stiffness and translational stiffness.

[0082] In S5, the set value of the total weighted error ≤ 0.01.

[0083] In S6, the calculated final value is used as the boundary equivalent stiffness to carry out the impact test analysis of the ground service equipment for the local panel 3, or as the stiffness value to be applied to the test piece of the local panel 3 to carry out the impact test of the ground service equipment.

[0084] It should be noted that the impact test analysis in the present invention specifically refers to the low - speed impact test analysis, and the low - speed impact generally refers to the impact with a speed range of 0 - 5 m / s.

[0085] The above are only the preferred embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent replacements, improvements, etc. made within the spirit and principles of the present invention shall be included within the protection scope of the present invention.

Claims

1. A method for determining the equivalent stiffness of the boundary of an aircraft panel structure, characterized in that, It includes the following steps: S1. Conduct finite element modeling on the fuselage to obtain a fuselage finite element model; apply a load at the first loading center point of the fuselage, gradually increase the load until the set load is reached, and obtain the first displacement-load curve and the first boundary rotation-load curve of the first loading center point; S2. Conduct finite element modeling on the local panel (3) to obtain a local panel (3) finite element model, where the local panel (3) is a partial arc plate of the fuselage; simplify the structure of the local panel (3), calculate and obtain the boundary stiffness of the local panel (3) as the initial boundary condition of the local panel (3) finite element model, apply a load at the second loading center point of the local panel (3), gradually increase the load until the set load is reached, and obtain the second displacement-load curve and the second boundary rotation-load curve of the second loading center point; S3. Compare the first displacement-load curve and the second displacement-load curve to obtain the first weighted error; compare the first boundary rotation-load curve and the second boundary rotation-load curve to obtain the second weighted error, and combine the first weighted error and the second weighted error to obtain the total weighted error; S4. Starting from the initial boundary condition, update the boundary stiffness through an optimization method, and calculate the second displacement-load curve and the second boundary rotation-load curve of the second loading center point based on the updated boundary stiffness; S5. Repeat S3 and S4 until the total weighted error is reduced to the set range, and the finally obtained boundary stiffness is the boundary equivalent stiffness; S6. Based on the boundary equivalent stiffness, conduct a ground service equipment impact test on the aircraft panel structure for the local panel (3).

2. The method for determining the equivalent stiffness of the boundary of an aircraft panel structure according to claim 1, characterized in that In S1, the fuselage is the whole fuselage (1) or a 1 / 4 fuselage structure (2).

3. A method for determining the equivalent stiffness of the boundary of an aircraft panel structure according to claim 1, characterized in that, In S1, when the fuselage is a 1 / 4 fuselage structure (2), the boundary condition of the fuselage finite element model is a symmetric boundary condition.

4. A method for determining the equivalent stiffness of the boundary of an aircraft panel structure according to claim 1, characterized in that, In S2, the local panel (3) is a panel structure including three frames (4) and five stringers (5), or a panel structure including three frames (4) and six stringers (5).

5. A method for determining the equivalent stiffness of the boundary of an aircraft panel structure according to claim 1, characterized in that In S1, the steps for conducting finite element modeling on the fuselage include the following: S101. Set the constitutive parameters and damage parameters of the fuselage material, set the ply direction, and obtain a fuselage model; S102. Conduct geometric simplification and shelling on the fuselage model; S103. Mesh the fuselage model after geometric simplification and shelling to obtain a meshed fuselage model; S104. Establish the skin (6), stringers (5), frames (4), and shear bands, as well as their connection relationships in the meshed fuselage model; S105. Apply boundary conditions to the fuselage model to obtain a fuselage finite element model.

6. The method for determining the equivalent stiffness of the boundary of an aircraft panel structure according to claim 1, wherein, In S2, the process of simplifying the structure of the local panel (3) is as follows: simplify the frames (4) in the local panel (3) into arcs, simplify the skin (6) into multiple arcs, the arc directions of the simplified skin (6) are the same as those of the simplified frames (4), the arc radian of the simplified skin (6) is the same as that of the simplified frames (4), and the stringers (5) and shear bands in the local panel (3) are ignored.

7. A method for determining the equivalent stiffness of the boundary of an aircraft panel structure according to claim 1, characterized in that In S2, the boundary stiffness includes rotational stiffness, longitudinal stiffness, and radial stiffness.

8. A method for determining the equivalent stiffness of the boundary of an aircraft panel structure according to claim 7, characterized in that In S1 and S2, the initial value of the applied load and the step size of the increasing load are equal.

9. A method for determining the equivalent stiffness of the boundary of an aircraft panel structure according to claim 1, characterized in that, In S3, the calculation formula for the total weighted error is as follows: wherein, is the total weighted error, and are the displacements of the loading points calculated from the finite element model of the fuselage and the finite element model of the local panel (3) respectively, and are the boundary rotation angles calculated from the finite element model of the fuselage and the finite element model of the local panel (3) respectively, and are the corresponding weighting coefficients.

10. The method for determining the equivalent stiffness of the boundary of an aircraft panel structure according to claim 1, wherein, In S5, the set range ≤ 0.01.

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