Method for determining equivalent stiffness of aircraft panel structure boundary

Abstract

The invention discloses an aircraft panel structure boundary equivalent stiffness determination method, and belongs to the technical field of aircraft structure strength. According to the method, finite element modeling is carried out on a fuselage and a local wall plate, loads are loaded at the loading center points of the two finite element models respectively, corresponding displacement load curves and boundary corner load curves are obtained, after the two displacement load curves and the two boundary corner load curves are compared, a total weighting error is obtained, and the total weighting error is calculated. By updating the boundary stiffness of the local wall plate in the finite element model, the total weighting error meets the requirement, and the finally obtained boundary stiffness serves as the boundary equivalent stiffness. According to the method, the deformation mode and the failure behavior of the local wallboard structure and the whole fuselage are consistent, a small-size local wallboard structure can be adopted to replace a large-size fuselage frame section structure, and the test cost can be remarkably reduced while the analysis efficiency is greatly improved.

Description

Technical Field

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

[0002] In the field of civil aviation, ground service equipment at airports plays a vital role. They include but are not limited to various types of airport ground handling vehicles, boarding bridges, gangways, maintenance stands, and baggage handling devices. These equipment play an indispensable role in the docking, maintenance, passenger boarding, and baggage loading of aircraft. However, during the docking phase of the aircraft, such ground service equipment may collide with the aircraft due to improper operation, equipment failure, or environmental factors. The accidental collision of such ground service equipment with the fuselage structure has become one of the main factors causing structural damage to civil aircraft, posing a serious threat to flight safety.

[0003] With the continuous advancement of aviation technology, the design of the new generation of aircraft adopts a large number of composite wall panel structures, which have performed well in reducing aircraft weight, improving fuel efficiency and enhancing the overall performance of the aircraft. However, when this composite wall panel structure is hit, its performance is completely different from that of traditional metal materials. Under the impact, the composite wall panel may not show obvious structural deformation, pits or cracks and other visually visible damage, but the impact may cause serious damage to the structure. These internal damages include but are not limited to the destruction of the fuselage frame, debonding of ribs or shear bands, shear failure of bolts, and fracture of ribs. These forms of damage not only seriously reduce the structural performance of the aircraft, but also may pose a direct threat to flight safety.

[0004] In view of the serious consequences that ground service equipment may bring when it hits an aircraft, in order to ensure the safety of aircraft service, it is necessary to conduct in-depth research on the relevant situations of ground service equipment hitting an aircraft. However, in 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 testing. Therefore, many studies have begun to focus on the simulation process, trying to reproduce the aircraft collision process through computer simulation to evaluate the impact of the collision on the aircraft structure.

[0005] Although simulation technology is increasingly used in the aviation field, it still faces many challenges when simulating the specific scenario of ground service equipment hitting an 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 panels during the actual impact process. This makes it difficult to ensure the accuracy of the simulation results, thus limiting the application of simulation technology in assessing aircraft structural damage.

[0006] On the other hand, if the whole machine is used for impact testing directly, although more accurate data can be obtained, it is obviously costly and difficult to achieve. Not only because the whole machine test consumes a lot of resources and funds, but also because of the safety risks and environmental impacts that may be brought about during the test. Summary of the invention

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

[0008] In order to achieve the above object, the present invention adopts the following technical solutions: A method for determining the equivalent stiffness of an aircraft panel structure boundary comprises the following steps: S1, performing finite element modeling on the fuselage to obtain a finite element model of the fuselage; applying a load to a first loading center point of the fuselage, gradually increasing the load until a set load is reached, and obtaining a first displacement load curve and a first boundary angle load curve of the first loading center point; S2, performing finite element modeling on the local wall panel to obtain a finite element model of the local wall panel; the local wall panel is a partially curved panel of the fuselage; simplifying the structure of the local wall panel, calculating and obtaining the boundary stiffness of the local wall panel as the initial boundary condition of the finite element model of the local wall panel, loading a load at a second loading center point of the local wall panel, gradually increasing the load until a set load is reached, and obtaining a second displacement load curve and a second boundary rotation angle load curve of the second loading center point; S3, comparing the first displacement load curve and the second displacement load curve to obtain a first weighted error; comparing the first boundary angle load curve and the second boundary angle load curve to obtain a second weighted error, and combining the first weighted error and the second weighted error to obtain a total weighted error; S4, starting from the initial boundary condition, updating the boundary stiffness by an optimization method, and calculating a second displacement load curve of a second loading center point and a second boundary rotation angle load curve based on the updated boundary stiffness; S5, repeat S3 and S4 until the total weighted error is reduced to a set range, and the boundary stiffness finally obtained is the boundary equivalent stiffness; S6, based on the boundary equivalent stiffness, the ground service equipment impact test of the aircraft panel structure is carried out for the local panel.

[0009] A further improvement of the present invention is: Preferably, in S1, the fuselage is a whole fuselage or a 1 / 4 fuselage structure.

[0010] 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.

[0011] Preferably, in S2, the local wall panel is a wall panel structure including three bulkheads and five long stringers, or a wall panel structure including three bulkheads and six long stringers.

[0012] Preferably, in S1, the performing finite element modeling on the fuselage comprises the following steps: S101, setting constitutive parameters and damage parameters of the fuselage material, setting the ply direction, and obtaining a fuselage model; S102, geometric simplification and shell extraction of the fuselage model; S103, meshing the fuselage model after geometric simplification and shell extraction to obtain a mesh fuselage model; S104, establish skin, long stringer, bulkhead and shear band in the mesh fuselage model, as well as their mutual connection relationships; S105, applying boundary conditions to the fuselage model to obtain a fuselage finite element model.

[0013] Preferably, in S2, the structural process of simplifying the local wall panel is: simplifying the bulkhead in the local wall panel into an arc, simplifying the skin into multiple arcs, the direction of the arc of the simplified skin is the same as the direction of the arc of the simplified bulkhead, the curvature of the arc of the simplified skin is the same as the curvature of the arc of the simplified bulkhead, and the long stringers and shear bands in the local wall panel are ignored.

[0014] Preferably, in S2, the boundary stiffness includes rotational stiffness, longitudinal stiffness and radial stiffness.

[0015] Preferably, in S1 and S2, the initial value of the applied load and the step length of increasing the load are equal.

[0016] Preferably, in S3, the calculation formula of the total weighted error is:

[0017] in, is the total weighted error, and are the loading point displacements calculated by the fuselage finite element model and the local panel finite element model, and are the boundary angles calculated by the fuselage finite element model and the local panel finite element model, and is the corresponding weighting coefficient.

[0018] Preferably, in S5, the setting range is ≤0.01.

[0019] Compared with the prior art, the present invention has the following beneficial effects: The invention discloses a method for determining the boundary equivalent stiffness of an aircraft wall panel structure. The method performs finite element modeling on a fuselage and a local wall panel respectively, applies loads at the loading center points in the two finite element models, obtains corresponding displacement load curves and boundary angle load curves respectively, obtains the first weighted errors of the two displacement load curves, and the second weighted errors of the two boundary angle load curves, obtains the total weighted error based on the first weighted error and the second weighted error, further updates the boundary stiffness of the local wall panel in the finite element model so that the total weighted error meets the requirements, and the boundary stiffness finally obtained is identified as the boundary equivalent stiffness. The method combines a simplified theoretical method and a fine finite element method, can accurately and quickly calculate the real boundary stiffness characteristics of the local wall panel of the aircraft through the finite element model, is applied to the local wall panel, and ensures that the boundary stiffness characteristics of the local wall panel are consistent with those when impacting the entire fuselage, so that the deformation mode and failure behavior of the local wall panel structure are consistent with those of the entire fuselage, so that the local wall panel structure can be used to carry out impact analysis and testing, and while reducing the analysis and test costs, the impact resistance characteristics analysis and test results of the fuselage composite wall panel structure are ensured to be authentic and reliable. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] Figure 1 A flow chart of a method for determining the equivalent stiffness of an aircraft panel structure boundary according to the present invention; Figure 2 It is a schematic diagram of the positional relationship between the overall fuselage and the local wall panel structure of the present invention; Figure 3 It is a schematic diagram of the positional relationship between the 1 / 4 fuselage structure and the local wall panel structure of the present invention; Figure 4 It is the main view of the local wall panel structure; Figure 5 It is a three-dimensional side view of the local wall panel structure; Figure 6 This is the finite element model diagram of the 1 / 4 fuselage structure; Figure 7 Schematic diagram of boundary stiffness simulation; Among them, 1. Integral fuselage; 2. 1 / 4 fuselage structure; 3. Partial wall panels; 4. Bulkheads; 5. Long spars; 6. Skin. DETAILED DESCRIPTION

[0021] In the following, the terms "first", "second", "third", and "fourth" are used for descriptive purposes only and should not be understood as indicating or implying relative importance or implicitly indicating the number of the indicated technical features. Therefore, a feature defined as "first", "second", "third", and "fourth" may explicitly or implicitly include one or more of the features.

[0022] It should be noted that the terms "first", "second", etc. in the specification and drawings of the present invention are used to distinguish similar objects, and are not necessarily used to describe a specific order or sequence. It should be understood that the data used in this way can be interchanged where appropriate, 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 "including" and "having" and any variations thereof are intended to cover non-exclusive inclusions, for example, a process, method, system, product or device that includes a series of steps or units is not necessarily limited to those steps or units that are clearly listed, but may include other steps or units that are not clearly listed or inherent to these processes, methods, products or devices.

[0023] The method of the present invention discloses a method for determining the boundary equivalent stiffness of an aircraft panel structure, comprising the following steps: S1, performing finite element modeling on the fuselage to obtain a finite element model of the fuselage; applying a load to a first loading center point of the fuselage, gradually increasing the load until a set load is reached, and obtaining a first displacement load curve and a first boundary angle load curve of the first loading center point; S2, perform finite element modeling on the local wall panel 3 to obtain a finite element model of the local wall panel 3; the local wall panel 3 is a partially curved panel on the fuselage; simplify the structure of the local wall panel 3, calculate and obtain the boundary stiffness of the local wall panel 3 as the initial boundary condition of the finite element model of the local wall panel 3, load the second loading center point of the local wall panel 3, gradually increase the load until the set load is reached, and obtain the second displacement load curve and the second boundary rotation angle load curve of the second loading center point; S3, comparing the first displacement load curve and the second displacement load curve to obtain a first weighted error; comparing the first boundary angle load curve and the second boundary angle load curve to obtain a second weighted error, and combining the first weighted error and the second weighted error to obtain a total weighted error; S4, starting from the initial boundary condition, updating the boundary stiffness by an optimization method, and calculating a second displacement load curve of a second loading center point and a second boundary rotation angle load curve based on the updated boundary stiffness; S5, repeat S3 and S4 until the total weighted error is reduced to a set range, and the boundary stiffness finally obtained is the boundary equivalent stiffness; S6, based on the boundary equivalent stiffness, for the local panel 3, a ground service equipment impact test of the aircraft panel structure is carried out.

[0024] The present invention adopts a local wall panel 3 to replace the overall fuselage 1 structure, and proposes an effective method and a feasible process for calculating the boundary stiffness of the local wall panel 3, so as to achieve the consistency of the deformation mode and failure behavior of the local wall panel 3 and the overall fuselage 1, thereby ensuring the reliability of the analysis or test results and significantly reducing the cost and difficulty of the analysis and test; the method combines the simplified theoretical method and the fine finite element method, and quickly calculates the estimated value of the boundary stiffness of the local wall panel 3 through the theoretical algorithm, and uses it as the initial value of the boundary stiffness of the finite element model of the local wall panel 3, and performs numerical iterative calculation to obtain the optimized boundary stiffness value, which not only improves the calculation efficiency but also ensures the accuracy of the stiffness calculation.

[0025] In some embodiments of the present invention, the fuselage in S1 is a whole fuselage 1 or a quarter fuselage structure 2, see Figure 2 This figure is a schematic diagram of the structural relationship between the overall fuselage 1 and the local wall panel 3. The actual length of the overall fuselage 1 is much longer than the length in the schematic diagram. Preferably, the local wall panel 3 is arranged at the center position of the overall fuselage 1 in the length direction.

[0026] See also Figure 3 , the 1 / 4 fuselage structure 2 is a curved surface, which is composed of a continuous array of arcs; the arc length is 1 / 4 of the circumference, the arc is an arc segment on the cross section perpendicular to the fuselage axis, and the arc corresponding to the arc is 90°. Preferably, the local wall panel 3 in the 1 / 4 fuselage structure 2 is still at the center position of the 1 / 4 fuselage structure 2 in the length direction.

[0027] pass Figure 2 and Figure 3 It can be seen that the local wall panel 3 is a part of the overall fuselage 1 or the 1 / 4 fuselage structure 2 and is a curved plate. The present invention limits the edge of the local wall panel 3 during the impact by using equivalent stiffness, which is equivalent to the local wall panel 3 still being in the overall fuselage 1 or the 1 / 4 fuselage structure 2 during the impact test, so that the impact test result is true and reliable.

[0028] Preferably, finite element modeling is performed on the 1 / 4 fuselage structure 2 to replace the entire fuselage 1, which can simplify the simulation process and speed up the solution efficiency.

[0029] In some embodiments of the present invention, see Figure 4 and Figure 5 In the overall fuselage 1 structure of the aircraft, the local effect is considered according to the selected impact area, and the structure and size of the local wall panel 3 are selected. An arc-shaped plate including three bulkheads 4 and five long stringers 5 is selected as the local wall panel 3, or an arc-shaped plate including three bulkheads 4 and six long stringers 5 is selected as the local wall panel 3; while ensuring the impact test effect, the model can be simplified.

[0030] See also Figure 4 and Figure 5, Figure 4 is the front view of the local wall panel 3, Figure 5 It is a three-dimensional side view of a partial wall panel 3, which includes a skin 6, a long stringer 5 and a bulkhead 4. The long stringer 5 is arranged along the axial direction of the entire fuselage, and the bulkhead 4 is arranged along the circumferential direction of the entire fuselage. The long stringer 5 and the bulkhead 4 are connected by a shear band, and the planes where the long stringer 5 and the bulkhead 4 are located are perpendicular to each other. It should be understood that the complete bulkhead 4 is a circular line, and what is shown in the figure is only a part.

[0031] In some embodiments of the present invention, in S1, the finite element modeling of the fuselage includes the following steps: S101, setting the constitutive parameters and damage parameters of the fuselage, setting the ply thickness and ply direction, and obtaining the fuselage model; S102, geometric simplification and shell extraction of the fuselage model; S103, meshing the fuselage model after geometric simplification and shell extraction, with quadrilateral meshes as the main mesh and triangular meshes as the auxiliary mesh, to obtain a mesh fuselage model; S104, establishing the skin 6, the long stringer 5, the bulkhead 4 and the shear band in the grid fuselage model, and the connection relationship between them; S105, applying boundary stiffness to the fuselage model. It should be noted that the boundary conditions applied to the entire fuselage 1 are boundary conditions applied on both sides. If the fuselage is a 1 / 4 fuselage structure 2, the boundary conditions applied are two symmetrical boundary conditions of the upper and lower and left and right. After the boundary conditions are applied, the finite element modeling of the entire fuselage is completed.

[0032] In some implementation schemes of the present invention, in S2, the finite element modeling process for the local wall panel 3 is the same as the finite element modeling process for the fuselage described above, but the boundary conditions applied are different.

[0033] For details, see Figure 6 In the process of finite element modeling of the local wall panel 3, the structure of the local wall panel 3 is simplified by using the simplified theory method according to its size and structural composition. The skin 6 and the bulkhead 4 can be simplified into one or more arcs with rectangular cross sections; the skin 6 is simplified into multiple arcs with rectangular cross sections and arrays, and the bulkhead 4 itself is simplified into an arc with a certain thickness and rectangular cross section; the direction of the arc of the simplified skin 6 is the same as that of the simplified bulkhead 4, and the arc of the simplified skin 6 is the same as that of the simplified bulkhead 4. Since the influence of the long stringer 5 and the shear band on the boundary stiffness is small, they can be ignored in the simulation calculation.

[0034] Preferably, see Figure 7 The boundary condition applied to the local wall plate 3 is the boundary stiffness, and the initial boundary condition applied is the calculated boundary stiffness, which is the inherent property value of the local wall plate 3. The boundary stiffness includes the rotational stiffness , longitudinal stiffness and radial stiffness ; The rotational stiffness is the bending stiffness of the local wall panel 3 boundary when subjected to horizontal radial load along the arc (the stiffness is represented by a spiral line in the figure); the longitudinal stiffness is the translation stiffness of the boundary perpendicular to the load direction when subjected to horizontal radial load along the arc; the radial stiffness is the translation stiffness of the boundary parallel to and opposite to the load direction when subjected to horizontal radial load along the arc. According to the actual situation during the collision, the rotational stiffness is decomposed into the stiffness of the upper and lower sides when subjected to force, the radial stiffness is decomposed into the stiffness of the upper and lower sides when subjected to force, and the longitudinal stiffness mainly relies on the support of the bottom, so only the radial stiffness constraint is set at the bottom. In the figure, F is the abbreviation of the force applied by the horizontal radial load.

[0035] The three boundary stiffness values ​​are calculated by the theoretical calculation formula of the three boundary stiffnesses, which are used as the initial boundary conditions for subsequent iterative calculations. , longitudinal stiffness and radial stiffness The calculation formulas are as follows: (1), (2) and (3):

[0036]

[0037]

[0038] in, is the elastic modulus, is Poisson's ratio, is the width of the arc section, is the equivalent thickness of the arc section, is the arc length, When simulating the local wall panel 3, the relevant parameters of the local wall panel 3 are brought in to calculate the boundary stiffness condition of the local wall panel 3 model. , and The initial value of .

[0039] 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. It can only be used as the input value of the subsequent parameter optimization automation process.

[0040] In some embodiments of the present invention, loading areas are set in the fuselage in S1 and the local wall panel 3 in S2. The loading areas are all actual impact areas and are part of the fuselage model or the local wall panel 3. In actual tests, the impact area refers to the area where the airport ground service equipment anti-collision strip contacts the skin 6, which is generally an elongated rectangular area, and its length generally does not exceed 2000mm and its width generally does not exceed 100mm. 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 wall panel 3. As a preferred solution, the two loading center points are both at the geometric center of the actual loading area.

[0041] In S1, the first loading center point is the actual impact point on the fuselage that is susceptible to impact or a simulated position, wherein the first displacement load curve represents a curve showing the relationship between the load change at the first loading center point and the position change of the first loading center point, and the first boundary angle load curve represents the relationship between the load change at the first loading center point and the boundary angle of the loading area, which boundary is the boundary of the loading area.

[0042] In some embodiments of the present invention, in S2, the second loading center point is the actual impact point of the local wall panel 3 during the test, and the second displacement load curve represents a curve showing the relationship between the load change at the second loading center point and the position change of the second loading center point, and the second boundary angle load curve represents the relationship between the load change at the second loading center point and the boundary angle of the loading area, which boundary is the boundary of the loading area.

[0043] It should be noted that in S1 and S2, the initial value of the loading load and the step size of the load increase are the same, so that during the simulation process, the local wall panel 3 and the fuselage are simulated under the load of the same working condition.

[0044] In some implementation schemes of the present invention, in S3, after the load is loaded to the set value, the loading point displacement and boundary rotation angle of the fuselage and the local wall panel 3 are compared, and two weighted errors, a first weighted error and a second weighted error, are calculated. The first weighted error and the second weighted error are respectively multiplied by corresponding coefficients and then added to obtain a total weighted error. At the beginning of the calculation, the total weighted error is difficult to meet the set range value, indicating that the loading stiffness of the local wall panel 3 is far from the simulation of the overall fuselage 1 or 1 / 4 fuselage structure 2. Therefore, the boundary stiffness value is updated based on the generalized gradient descent method, and iterative calculation is performed.

[0045] The weighted error calculation formula is: (4) in, and are the loading point displacements calculated by the fuselage finite element model and the local panel 3 finite element model, and are the boundary angles calculated by the fuselage finite element model and the local panel 3 finite element model, and is the corresponding weighting coefficient, and The value range of is 0-1, and + =1, the default value range is = =0.5, and can be adjusted according to the requirements for rotational stiffness and translational stiffness.

[0046] In S5, the setting value of the total weighted error is ≤ 0.01.

[0047] In S6, the calculated final value is used as the boundary equivalent stiffness to carry out the ground service equipment impact test analysis of the local wall panel 3, or is used as the stiffness value applied to the local wall panel 3 test piece to carry out the ground service equipment impact test.

[0048] It should be noted that the impact test analysis in the present invention specifically refers to low-speed impact test analysis, and low-speed impact generally refers to impact with a speed range of 0-5 m / s.

[0049] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principle of the present invention should be included in 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: The following steps are involved: S1, performing finite element modeling on the fuselage to obtain a finite element model of the fuselage; applying a load to a first loading center point of the fuselage, gradually increasing the load until a set load is reached, and obtaining a first displacement load curve and a first boundary angle load curve of the first loading center point; S2, performing finite element modeling on the local wall panel (3) to obtain a finite element model of the local wall panel (3), wherein the local wall panel (3) is a partially curved panel of the fuselage; simplifying the structure of the local wall panel (3), calculating and obtaining the boundary stiffness of the local wall panel (3) as an initial boundary condition of the finite element model of the local wall panel (3), applying a load to a second loading center point of the local wall panel (3), gradually increasing the load until a set load is reached, and obtaining a second displacement load curve and a second boundary rotation angle load curve of the second loading center point; S3, comparing the first displacement load curve and the second displacement load curve to obtain a first weighted error; comparing the first boundary angle load curve and the second boundary angle load curve to obtain a second weighted error, and combining the first weighted error and the second weighted error to obtain a total weighted error; S4, starting from the initial boundary condition, updating the boundary stiffness by an optimization method, and calculating a second displacement load curve of a second loading center point and a second boundary rotation angle load curve based on the updated boundary stiffness; S5, repeat S3 and S4 until the total weighted error is reduced to a set range, and the boundary stiffness finally obtained is the boundary equivalent stiffness; S6, based on the boundary equivalent stiffness, a ground service equipment impact test of the aircraft panel structure is carried out for the local panel (3).

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

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

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

5. 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 finite element modeling of the fuselage includes the following steps: S101, setting constitutive parameters and damage parameters of the fuselage material, setting the ply direction, and obtaining a fuselage model; S102, geometric simplification and shell extraction of the fuselage model; S103, meshing the fuselage model after geometric simplification and shell extraction to obtain a mesh fuselage model; S104, establishing the skin (6), the long stringer (5), the bulkhead (4) and the shear band in the mesh fuselage model, and the connection relationship between them; S105, applying boundary conditions to the fuselage model to obtain a fuselage finite element model.

6. A method for determining the equivalent stiffness of an aircraft panel structure boundary according to claim 1, characterized in that: In S2, the structural process of simplifying the local wall panel (3) is as follows: simplifying the bulkhead (4) in the local wall panel (3) into an arc, simplifying the skin (6) into a plurality of arcs, the direction of the simplified arc of the skin (6) being the same as the direction of the simplified arc of the bulkhead (4), the arc angle of the simplified arc of the skin (6) being the same as the arc angle of the simplified arc of the bulkhead (4), and the long girder (5) and the shear band in the local wall panel (3) being ignored.

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

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

10. The method for determining the equivalent boundary stiffness of an aircraft panel structure according to claim 1, characterized in that: In S5, the setting range is ≤0.01.

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