Method for predicting assembly deformation of equal straight section of fuselage in fixed-wing aircraft

Abstract

The invention provides a fixed-wing aircraft fuselage equal straight section assembly deformation prediction method, which comprises the following steps: establishing a three-dimensional simplified structure model of a skin assembly, a fuselage frame assembly and a bottom plate beam assembly according to geometric features; the skin assembly, the fuselage frame assembly and the bottom plate beam assembly are assembled, a constraint relation is applied, and a three-dimensional simplified structure model of the straight section of the middle fuselage is formed; grid division is carried out on the three-dimensional simplified structure model of the straight section of the middle fuselage; setting material attributes and mechanical boundary conditions for the three-dimensional simplified structure model of the straight section of the middle fuselage; solving a deformation result of the three-dimensional simplified structure model for the straight section of the middle fuselage; and evaluating and optimizing a calculation result. According to the method, the influence of the skin size of the middle fuselage straight section, the fuselage frame assembly size, the bottom plate beam size, the material attribute and the connection mode can be considered, and the deformation of the middle fuselage straight section of the fixed-wing aircraft can be rapidly calculated.

Description

Technical Field

[0001] The invention belongs to the technical field of aircraft component assembly, and in particular relates to a method for predicting assembly deformation of straight sections such as a fuselage in a fixed-wing aircraft. Background Art

[0002] The mid-fuselage of a fixed-wing aircraft serves as the core load-bearing structure connecting the wings, displacement, and payload compartments. Its straight sections have the geometric characteristics of large, thin-walled cylinders, meeting the requirements of high strength, lightweight, and aerodynamic shape accuracy. As aircraft become larger, the size of straight sections of the mid-fuselage increases. At the same time, due to structural weight restrictions, the weak rigidity of the structure becomes increasingly apparent. Traditional experience-based assembly process compensation is no longer able to meet the structural characteristics of large, weakly rigid mid-fuselage sections in aircraft. It is necessary to establish a method for rapidly predicting the assembly deformation of straight sections of the mid-fuselage of fixed-wing aircraft, so as to further design process solutions and assembly tooling solutions.

[0003] For straight sections like the fuselage in fixed-wing aircraft, modeling and analyzing their assembly deformation is difficult due to the hundreds of thousands of parts required. Consequently, there is a lack of methods for predicting the assembly deformation of these straight sections. Traditional methods simplify the main components into two-dimensional beam and shell elements for calculation and deformation prediction. However, such simplification results in a loss of local detail in the model and a failure to account for the deformation behavior of the true three-dimensional structure, significantly deviating from actual conditions. This makes it impossible to effectively capture the assembly deformation of straight sections like the fuselage in fixed-wing aircraft, leading to inappropriate assembly plans and tooling design. Summary of the Invention

[0004] The purpose of the present invention is: The embodiment of the present invention proposes a method for predicting the assembly deformation of straight sections such as the fuselage in fixed-wing aircraft, so as to solve the problem that traditional calculation models lose local detail features, cannot consider three-dimensional deformation effects, cannot obtain microscopic deformation results in the thickness direction, and are significantly different from actual conditions, resulting in the inability to accurately obtain the assembly deformation of straight sections such as the fuselage in fixed-wing aircraft, leading to unreasonable assembly plan formulation and assembly tooling design, resulting in reduced assembly quality.

[0005] The present application provides a method for predicting the assembly deformation of straight sections such as the fuselage in a fixed-wing aircraft, the method comprising the following steps:

[0006] Step 1: Based on the geometric features, a three-dimensional simplified structural model of the skin component, a three-dimensional simplified structural model of the fuselage frame component, and a three-dimensional simplified structural model of the bottom beam component are established;

[0007] Step 2: Assemble the 3D simplified structural model of the skin assembly, the 3D simplified structural model of the fuselage frame assembly, and the 3D simplified structural model of the bottom plate beam assembly, and apply constraints to form a 3D simplified structural model of the straight section such as the mid-fuselage;

[0008] Step 3: Mesh the simplified 3D structural model of the straight sections such as the mid-fuselage.

[0009] Step 4: Set material properties and mechanical boundary conditions for the simplified 3D structural model of the mid-fuselage and other straight sections;

[0010] Step 5: Calculate the deformation results of the simplified three-dimensional structural model for straight sections such as the mid-fuselage;

[0011] Step 6: Evaluation and optimization of calculation results.

[0012] Preferably, the step 1 includes:

[0013] Step 1-1, using CAD software to establish a three-dimensional simplified structural model of the mid-fuselage straight section skin component by using the mid-fuselage straight section radius, mid-fuselage straight section length, and skin thickness;

[0014] Step 1-2, establishing a simplified three-dimensional structural model of the fuselage frame assembly using the fuselage frame outer diameter and inner diameter, frame thickness, and frame length;

[0015] Steps 1-3: establishing a simplified three-dimensional structural model of the floor beam assembly by using the floor beam assembly length, floor beam width, floor beam height, and floor beam cross-sectional geometry;

[0016] Preferably, the simplified models of the straight sections such as the mid-fuselage obtained in steps 1-1 to 1-3 have the following requirements:

[0017] The number and spacing of the three-dimensional simplified structural models of the skin components, the three-dimensional simplified structural models of the fuselage frame components, and the three-dimensional simplified structural models of the bottom plate beam components are required to be consistent with the actual components, and the length, width, and height dimensions of the three-dimensional simplified structural models of the skin components, the three-dimensional simplified structural models of the fuselage frame components, and the three-dimensional simplified structural models of the bottom plate beam components are required to differ by less than 5% from the actual components.

[0018] Preferably, the step 2 includes:

[0019] Step 2-1, assembling the three-dimensional simplified structural model of the skin assembly, the three-dimensional simplified structural model of the fuselage frame assembly, and the three-dimensional simplified structural model of the bottom beam assembly according to the assembly process plan;

[0020] Step 2-2: The constraint equation between the inner surface of the 3D simplified structural model of the skin component and the outer surface of the 3D simplified structural model of the fuselage frame component is:

[0021]

[0022] Among them, u1, u2, and u3 represent the three displacement components of the aircraft in the heading, span, and gravity directions respectively; the superscripts M and S represent the inner surface of the three-dimensional simplified structural model of the skin component and the outer surface of the three-dimensional simplified structural model of the fuselage frame component respectively;

[0023] In step 2-3, the constraint equation between the inner surface of the 3D simplified structural model of the skin component and the outer surface of the 3D simplified structural model of the bottom beam component is:

[0024]

[0025] Among them, u1, u2, and u3 represent the three displacement components of the aircraft in the heading, span, and gravity directions, respectively; the superscripts M and V represent the inner surface of the three-dimensional simplified structural model of the skin component and the outer surface of the three-dimensional simplified structural model of the bottom plate beam component, respectively, forming a three-dimensional simplified structural model of straight sections such as the mid-fuselage.

[0026] Preferably, the step three has the following requirements:

[0027] The three-dimensional simplified structural model of the straight section such as the mid-fuselage is divided into grid units. The maximum length, width and height of the grid units in the three-dimensional simplified structural model of the straight section such as the mid-fuselage shall not exceed 10% of the length, width and height of each component. It is required to check the geometric dimensions of all grid units. If the requirement exceeds 10%, the grid size needs to be modified and the grid division needs to be re-performed until the requirements are met.

[0028] Preferably, the step 4 includes:

[0029] Step 2-1: Set the Young's modulus, Poisson's ratio, and material density of the simplified model based on the material types of the actual skin component, fuselage frame component, and bottom beam component.

[0030] Step 2-2: setting 6-DOF constraints on the two end faces of the 3D simplified structural model of the mid-fuselage straight section as mechanical boundary conditions of the 3D simplified structural model of the mid-fuselage straight section.

[0031] Preferably, the step five includes:

[0032] Step 5-1, apply gravity load to the simplified model of the entire straight section of the middle fuselage;

[0033] Step 5-2, solve the problem using the statics solution method to calculate the deformation results of the simplified straight section model such as the fuselage under gravity load.

[0034] Preferably, the step six includes:

[0035] In step 6-1, replace the boundary conditions of the simplified model of the entire mid-fuselage and other straight sections with the actual location of the support structure at the assembly site, and repeat step 5 to obtain the deformation results under the actual assembly site;

[0036] Step 6-2: Use the actual measurement results at the assembly site as a standard and compare them with the calculated deformation results. If the relative error between the two is less than 10%, the calculation is considered qualified.

[0037] In step 6-3, if the relative error between the calculated deformation result and the actual measurement result at the assembly site is greater than 10%, the thickness value in the 3D simplified structural model of each component is reduced, and steps 2 to 6 are repeated until the result meets the requirements; if the relative error between the calculated deformation result and the actual measurement result at the assembly site is less than 10%, the thickness value in the 3D simplified structural model of each component is increased, and steps 2 to 6 are repeated until the result meets the requirements.

[0038] The beneficial effects of the present invention are:

[0039] The present invention aims to solve the problem that traditional calculation models lose local detail features, cannot consider three-dimensional deformation effects, cannot obtain microscopic deformation results in the thickness direction, and are significantly different from actual conditions, resulting in the inability to accurately obtain the assembly deformation of straight sections such as the fuselage in fixed-wing aircraft, leading to unreasonable assembly plan formulation and assembly tooling design, resulting in reduced assembly quality. The method can take into account the influence of the skin size of straight sections such as the fuselage, the size of the fuselage frame components, the size of the bottom plate beam, the material properties, and the connection method, and quickly calculate the deformation of straight sections such as the fuselage in fixed-wing aircraft. The established rapid calculation method can provide a theoretical basis for the determination of assembly plans and the design of assembly tooling, further improving the assembly quality of straight sections such as the fuselage in fixed-wing aircraft, and is a fast and effective calculation method. BRIEF DESCRIPTION OF THE DRAWINGS

[0040] Figure 1 This is a flow chart of a method for predicting assembly deformation of straight sections such as the fuselage in a fixed-wing aircraft proposed by the present invention;

[0041] Figure 2 The three-dimensional simplified structural model of the skin assembly, the three-dimensional simplified structural model of the fuselage frame assembly, and the three-dimensional simplified structural model of the bottom beam assembly in step 1 of the present invention;

[0042] Figure 3 Meshing the simplified three-dimensional structural model of the straight sections such as the fuselage in step 3 of the present invention;

[0043] Figure 4 The skin deformation result is calculated in step 4 of the present invention;

[0044] Figure 5 Comparison of calculation results of different simplified models proposed in this invention;

[0045] Explanation of the numbers in the figure: 1. Three-dimensional simplified structural model of the skin component; 2. Three-dimensional simplified structural model of the fuselage frame component; 3. Three-dimensional simplified structural model of the bottom plate beam component; 4. Mesh division of the three-dimensional simplified structural model of the skin component; 5. Mesh division of the three-dimensional simplified structural model of the fuselage frame component; 6. Mesh division of the three-dimensional simplified structural model of the bottom plate beam component. DETAILED DESCRIPTION

[0046] The aforementioned background technology has demonstrated the importance of rapid calculation of assembly deformation for straight sections, such as the fuselage, in fixed-wing aircraft. However, traditional calculation models lack local detail features, fail to account for three-dimensional deformation effects, and are unable to obtain microscopic deformation results in the thickness direction, resulting in significant discrepancies with actual conditions. This results in an inability to accurately determine assembly deformation for straight sections, such as the fuselage, in fixed-wing aircraft. This leads to irrational assembly plan formulation and assembly tooling design, resulting in reduced assembly quality. To address these issues, embodiments of the present invention propose a method for predicting assembly deformation for straight sections, such as the fuselage, in fixed-wing aircraft. This method rapidly calculates deformation for straight sections, such as the fuselage, by considering the effects of skin dimensions, fuselage frame assembly dimensions, floor beam dimensions, material properties, and connection methods. The established rapid calculation method provides a theoretical basis for determining assembly plans and designing assembly tooling, further improving the assembly quality of straight sections, such as the fuselage, in fixed-wing aircraft. The present invention is further described in detail below with reference to the accompanying drawings and specific embodiments.

[0047] like Figure 1-Figure 5 As shown, the present application provides a method for predicting the assembly deformation of straight sections such as the fuselage in a fixed-wing aircraft, comprising the following steps:

[0048] Step 1: According to the geometric features, a three-dimensional simplified structural model of the skin component, a three-dimensional simplified structural model of the fuselage frame component, and a three-dimensional simplified structural model of the bottom beam component are established, such as Figure 2 As shown;

[0049] Among them, the three-dimensional simplified structural model of the skin component, the three-dimensional simplified structural model of the fuselage frame component, and the three-dimensional simplified structural model of the bottom plate beam component are requirements put forward considering the rigidity characteristics of straight section structures such as the mid-fuselage. Figure 5 The comparison results of the three-dimensional simplified structural model considering only the skin component, the three-dimensional simplified structural model considering the skin component and the fuselage frame component, and the three-dimensional simplified structural model considering the skin component, the fuselage frame component, and the bottom plate beam component are given. Figure 5It can be seen that the deformation prediction results of the three-dimensional simplified structural model considering only the skin component show fluctuating behavior, indicating that this is an abnormal behavior caused by an inappropriate simplified model. Considering the three-dimensional simplified structural models of the skin component and the fuselage frame component, the deformation trend is consistent with the actual situation, but the maximum deformation value is still far greater than the actual measurement size on site, which is inconsistent with the actual situation. Considering the three-dimensional simplified structural models of the skin component, the fuselage frame component, and the bottom plate beam component, not only the deformation trend is consistent with the actual situation, but also the maximum deformation value is close to the actual value. It should be pointed out that as the model volume increases, the calculation efficiency decreases. In order to take into account both calculation efficiency and calculation accuracy, the prediction model can include the three-dimensional simplified structure of the skin component, the fuselage frame component, and the bottom plate beam component. The step one includes:

[0050] Step 1-1: In ABAQUS 6.13 software, a simplified three-dimensional structural model of the mid-fuselage straight section skin component is established by setting the mid-fuselage straight section radius to 4000 mm, the mid-fuselage straight section length to 15000 mm, and the skin thickness to 2 mm. Figure 2 As shown;

[0051] Step 1-2, by taking the outer diameter of the fuselage frame as 4000mm and the inner diameter as 3900, the frame thickness as 20mm, and the frame length, a three-dimensional simplified structural model of the fuselage frame assembly is established, such as Figure 2 As shown;

[0052] Steps 1-3: Establish a simplified three-dimensional structural model of the bottom plate beam assembly by using the bottom plate beam assembly length, bottom plate beam width, bottom plate beam height, and bottom plate beam cross-sectional geometry, such as Figure 2 As shown;

[0053] The simplified models of straight sections such as the mid-fuselage obtained in steps 1-1 to 1-3 have the following requirements:

[0054] The number and spacing of the three-dimensional simplified structural models of the skin components, the three-dimensional simplified structural models of the fuselage frame components, and the three-dimensional simplified structural models of the bottom plate beam components are required to be consistent with the actual components, and the length, width, and height dimensions of the three-dimensional simplified structural models of the skin components, the three-dimensional simplified structural models of the fuselage frame components, and the three-dimensional simplified structural models of the bottom plate beam components are required to differ by less than 5% from the actual components.

[0055] Among them, through Figure 5It was found that the number and spacing of skin components, fuselage frame components, and bottom plate beam components have a significant impact on assembly deformation, and it is necessary to ensure that the number and spacing of each component are consistent with the actual object. Taking into account the influence of part processing accuracy on the actual skin components, fuselage frame components, and bottom plate beam components, there is a certain error between the physical model and the theoretical model. In addition, in order to improve the efficiency of mesh unit division of skin curvature, the outer contour detail size of the skin component can be adjusted. In this method, in order to improve the calculation efficiency, the influence of part processing errors is ignored, and the difference between the length, width, and height dimensions and the actual components is controlled within 5%.

[0056] Step 2: Assemble the three-dimensional simplified structural model of the skin assembly, the three-dimensional simplified structural model of the fuselage frame assembly, and the three-dimensional simplified structural model of the bottom plate beam assembly, apply constraint equations, and form a three-dimensional simplified structural model of the straight section such as the mid-fuselage;

[0057] The second step includes:

[0058] Step 2-1, assembling the three-dimensional simplified structural model of the skin assembly, the three-dimensional simplified structural model of the fuselage frame assembly, and the three-dimensional simplified structural model of the bottom beam assembly according to the assembly process plan;

[0059] Step 2-2: The constraint equation between the inner surface of the 3D simplified structural model of the skin component and the outer surface of the 3D simplified structural model of the fuselage frame component is:

[0060]

[0061] Among them, u1, u2, and u3 represent the three displacement components of the aircraft in the heading, span, and gravity directions respectively; the superscripts M and S represent the inner surface of the three-dimensional simplified structural model of the skin component and the outer surface of the three-dimensional simplified structural model of the fuselage frame component respectively;

[0062] In step 2-3, the constraint equation between the inner surface of the 3D simplified structural model of the skin component and the outer surface of the 3D simplified structural model of the bottom beam component is:

[0063]

[0064] Among them, u1, u2, and u3 represent the three displacement components of the aircraft in the heading, span, and gravity directions, respectively; the superscripts M and V represent the inner surface of the three-dimensional simplified structural model of the skin component and the outer surface of the three-dimensional simplified structural model of the bottom plate beam component, respectively, forming a three-dimensional simplified structural model of straight sections such as the mid-fuselage.

[0065] Step 3: Mesh the simplified 3D structural model of the straight sections such as the mid-fuselage.

[0066] Step 3 has the following requirements:

[0067] Divide the three-dimensional simplified structural model of the straight section such as the middle fuselage into grid units. The maximum length, width and height of the grid units in the three-dimensional simplified structural model of the straight section such as the middle fuselage shall not exceed 10% of the length, width and height of each component. It is required to check the geometric dimensions of all grid units. If the requirements exceed 10%, the grid size needs to be modified and the grid division needs to be re-performed until the requirements are met.

[0068] Step 4: Set material properties and mechanical boundary conditions for the simplified 3D structural model of the mid-fuselage and other straight sections;

[0069] The fourth step includes:

[0070] Step 2-1: Set the Young's modulus, Poisson's ratio, and material density of the simplified model according to the material types of the actual skin component, fuselage frame component, and bottom beam component. In this embodiment, they are all set to 70 GPa and 0.30, and the material density is 2.1×10 -9 t / mm 3 ;

[0071] Step 2-2: setting 6-DOF constraints on the two end faces of the 3D simplified structural model of the mid-fuselage straight section as mechanical boundary conditions of the 3D simplified structural model of the mid-fuselage straight section.

[0072] Step 5, solving the deformation result of the three-dimensional simplified structural model for the straight section such as the mid-fuselage;

[0073] The step five includes:

[0074] Step 5-1: Apply gravity load to the simplified model of the entire straight section of the middle fuselage, taking the gravity acceleration as 9800 mm / s 2 ;

[0075] Step 5-2, solve by statics method, and calculate the deformation results of the simplified straight section model of the fuselage under gravity load, such as Figure 4 shown.

[0076] Step 6: Evaluation and optimization of calculation results.

[0077] The step six comprises:

[0078] In step 6-1, the boundary conditions of the simplified model of the entire mid-fuselage and other straight sections are replaced with the actual location of the support structure at the assembly site. Step 5 is repeated to obtain the deformation results under the actual assembly site. In this embodiment, the maximum deformation of the skin is 1.15 mm, which is 1.1 mm when combined with the on-site measurement result. The difference between the on-site measurement result and the calculation method is less than 10%, indicating that the calculation method meets the requirements.

[0079] The above is a further detailed description of the present invention in conjunction with specific preferred embodiments, and the specific implementation of the present invention should not be considered to be limited to these descriptions. For those skilled in the art of the present invention, without departing from the concept of the present invention, several simple deductions or substitutions can be made, which should be considered to fall within the scope of protection of the present invention.

Claims

1. A method for predicting the assembly deformation of straight sections such as the fuselage of a fixed-wing aircraft, characterized in that: The method comprises the following steps: Step 1: Based on the geometric features, a three-dimensional simplified structural model of the skin component, a three-dimensional simplified structural model of the fuselage frame component, and a three-dimensional simplified structural model of the bottom beam component are established; Step 2: Assemble the 3D simplified structural model of the skin assembly, the 3D simplified structural model of the fuselage frame assembly, and the 3D simplified structural model of the bottom plate beam assembly, and apply constraints to form a 3D simplified structural model of the straight section such as the mid-fuselage; Step 3: Mesh the simplified 3D structural model of the straight sections such as the mid-fuselage. Step 4: Set material properties and mechanical boundary conditions for the simplified 3D structural model of the mid-fuselage and other straight sections; Step 5: Calculate the deformation results of the simplified three-dimensional structural model for straight sections such as the mid-fuselage; Step 6: Evaluation and optimization of calculation results.

2. The method for predicting assembly deformation of straight sections such as fuselage in fixed-wing aircraft according to claim 1, characterized in that: The step one comprises: Step 1-1, using CAD software to establish a three-dimensional simplified structural model of the mid-fuselage straight section skin component by using the mid-fuselage straight section radius, mid-fuselage straight section length, and skin thickness; Step 1-2, establishing a simplified three-dimensional structural model of the fuselage frame assembly using the fuselage frame outer diameter and inner diameter, frame thickness, and frame length; Steps 1-3: Establish a three-dimensional simplified structural model of the bottom plate beam assembly through the bottom plate beam assembly length, bottom plate beam width, bottom plate beam height, and bottom plate beam cross-sectional geometry.

3. The method for predicting assembly deformation of straight sections such as fuselage in fixed-wing aircraft according to claim 2, characterized in that: The simplified model of the straight section such as the mid-fuselage obtained in steps 1-1 to 1-3 has the following requirements: The number and spacing of the three-dimensional simplified structural models of the skin components, the three-dimensional simplified structural models of the fuselage frame components, and the three-dimensional simplified structural models of the bottom plate beam components are required to be consistent with the actual components, and the length, width, and height dimensions of the three-dimensional simplified structural models of the skin components, the three-dimensional simplified structural models of the fuselage frame components, and the three-dimensional simplified structural models of the bottom plate beam components are required to differ by less than 5% from the actual components.

4. The method for predicting assembly deformation of straight sections such as fuselage in fixed-wing aircraft according to claim 1, characterized in that: The second step includes: Step 2-1, assembling the three-dimensional simplified structural model of the skin assembly, the three-dimensional simplified structural model of the fuselage frame assembly, and the three-dimensional simplified structural model of the bottom beam assembly according to the assembly process plan; Step 2-2: The constraint equation between the inner surface of the 3D simplified structural model of the skin component and the outer surface of the 3D simplified structural model of the fuselage frame component is: Among them, u1, u2, and u3 represent the three displacement components of the aircraft in the heading, span, and gravity directions respectively; the superscripts M and S represent the inner surface of the three-dimensional simplified structural model of the skin component and the outer surface of the three-dimensional simplified structural model of the fuselage frame component respectively; In step 2-3, the constraint equation between the inner surface of the 3D simplified structural model of the skin component and the outer surface of the 3D simplified structural model of the bottom beam component is: Among them, u1, u2, and u3 represent the three displacement components of the aircraft in the heading, span, and gravity directions, respectively; the superscripts M and V represent the inner surface of the three-dimensional simplified structural model of the skin component and the outer surface of the three-dimensional simplified structural model of the bottom plate beam component, respectively, forming a three-dimensional simplified structural model of straight sections such as the mid-fuselage.

5. The method for predicting assembly deformation of straight sections such as fuselage in fixed-wing aircraft according to claim 2, characterized in that: Step 3 has the following requirements: The three-dimensional simplified structural model of the straight section such as the mid-fuselage is divided into grid units. The maximum length, width and height of the grid units in the three-dimensional simplified structural model of the straight section such as the mid-fuselage shall not exceed 10% of the length, width and height of each component. It is required to check the geometric dimensions of all grid units. If the requirement exceeds 10%, the grid size needs to be modified and the grid division needs to be re-performed until the requirements are met.

6. The method for predicting assembly deformation of straight sections such as fuselage in fixed-wing aircraft according to claim 4, characterized in that: The fourth step includes: Step 2-1: Set the Young's modulus, Poisson's ratio, and material density of the simplified model based on the material types of the actual skin component, fuselage frame component, and bottom beam component. Step 2-2: setting 6-DOF constraints on the two end faces of the 3D simplified structural model of the mid-fuselage straight section as mechanical boundary conditions of the 3D simplified structural model of the mid-fuselage straight section.

7. The method for predicting assembly deformation of straight sections such as fuselage in fixed-wing aircraft according to claim 1, characterized in that: The step five includes: Step 5-1, apply gravity load to the simplified model of the entire straight section of the middle fuselage; Step 5-2, solve the problem using the statics solution method to calculate the deformation results of the simplified straight section model such as the fuselage under gravity load.

8. The method for predicting assembly deformation of straight sections such as fuselage in fixed-wing aircraft according to claim 1, characterized in that: The step six comprises: In step 6-1, replace the boundary conditions of the simplified model of the entire mid-fuselage and other straight sections with the actual location of the support structure at the assembly site, and repeat step 5 to obtain the deformation results under the actual assembly site; Step 6-2: Use the actual measurement results at the assembly site as a standard and compare them with the calculated deformation results. If the relative error between the two is less than 10%, the calculation is considered qualified. In step 6-3, if the relative error between the calculated deformation result and the actual measurement result at the assembly site is greater than 10%, the thickness value in the 3D simplified structural model of each component is reduced, and steps 2 to 6 are repeated until the result meets the requirements; if the relative error between the calculated deformation result and the actual measurement result at the assembly site is less than 10%, the thickness value in the 3D simplified structural model of each component is increased, and steps 2 to 6 are repeated until the result meets the requirements.

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