Method and device for acquiring convergence stress of aircraft structure body

Abstract

The invention provides a method and device for acquiring convergence stress of an aircraft structure body, and the method comprises the steps: taking the minimum thickness t of the aircraft structure body as a reference size, and building an aircraft structure body finite element model through finite element analysis software; obtaining a maximum stress calculation result of the aircraft structure body according to the boundary condition and the load working condition, and recording the maximum stress calculation result as sigma1; the value of n is 2; taking 1 / n of the minimum thickness t of the aircraft structure body as a reference size, establishing an aircraft structure body finite element model, and obtaining a maximum stress calculation result of the aircraft structure body taking 1 / n of the minimum thickness t as the reference size according to a boundary condition and a load working condition, and recording the maximum stress calculation result as sigma n; judging whether (sigman-sigman-1) / sigman-1 is less than or equal to 0.02 or not, if not, adding 1 to the value of n, and calculating again until the condition is met; and fitting a convergence stress function according to n calculation data from sigma 1 to sigma n after the condition is met, and taking a limit value of the convergence stress function as the convergence stress of the aircraft structure body. According to the invention, the stress of the structure body can be accurately calculated.

Description

Technical Field

[0001] The present invention belongs to the field of aircraft structure strength design, and in particular relates to a method and device for obtaining convergent stress of an aircraft structure body. Background Art

[0002] Aircraft structures typically include the fuselage, wings, tailplane, engine nacelles, landing gear, control systems, and other structural components. The primary load-bearing components of an aircraft structure are often referred to as structural bodies, such as the wing bodies and the wing-to-fuselage joint bodies. The fundamental task of aircraft structural strength design is to design an airframe that meets operational requirements, meets acceptable strength, stiffness, fatigue, and damage tolerance, exhibits good manufacturability, and meets weight specifications, based on the aircraft model's design technical requirements and incorporating them into the basic structural design requirements.

[0003] However, structural body stress calculation and strength verification are often encountered in the field of aircraft structural strength design. The accuracy of structural body strength calculation is not only inevitably related to structural weight reduction and discovery of design weaknesses during the design phase, but also related to the flight safety of the aircraft. It provides theoretical support and guidance for structural strength design and daily aircraft inspection and maintenance, and is the basic data support for aircraft to meet economic and maintainability requirements.

[0004] Currently, structural strength analysis and calculations typically follow the principles of geometric and physical approximation, simplifying the structure into densely meshed plate or volume elements. Stress analysis is then performed directly, and the results from a single analysis are used directly for strength verification. This calculated result often deviates significantly from the maximum converged stress of the actual structure, posing a significant risk to aircraft flight safety. Therefore, a method for calculating the convergent stress of structural strength is needed to accurately calculate the stress magnitude of the structural body.

[0005] For aircraft structural elements, the accuracy of structural strength calculations is directly related to the flight safety, economy, and maintainability of the aircraft. Therefore, it is necessary to establish and propose a simple calculation method to conveniently, quickly, and accurately calculate the convergence stress of aircraft structural elements. Summary of the Invention

[0006] The present invention provides a method and device for obtaining the convergence stress of an aircraft structure body, and accurately calculates the stress magnitude of the structure body.

[0007] A first aspect of the present invention provides a method for obtaining convergent stress of an aircraft structure, which mainly includes the following steps:

[0008] S1. Using the minimum thickness t of the aircraft structure as a reference dimension, a finite element model of the aircraft structure is established using finite element analysis software;

[0009] S2. According to the boundary conditions and load conditions, the maximum stress calculation result of the aircraft structure body is obtained, which is recorded as σ1; let the value of n be 2;

[0010] S3. Taking 1 / n of the minimum thickness t of the aircraft structure as the reference size, the finite element model of the aircraft structure is established using finite element analysis software. According to the boundary conditions and load conditions, the maximum stress calculation result of the aircraft structure with 1 / n of the minimum thickness t as the reference size is obtained and recorded as σ n ;

[0011] S4, judge whether it satisfies (σ n -σ n-1 ) / σ n-1 ≤0.02. If not satisfied, add 1 to the value of n and return to S3. If satisfied, proceed to S5.

[0012] S5, according to σ1 to σ n The convergent stress function is fitted based on n calculation data, and the limit value of the convergent stress function is used as the convergent stress of the aircraft structure.

[0013] Optionally, when the final value of n is 2, the convergence stress function is convergence stress = σ n .

[0014] Optionally, when the final value of n is greater than 2, the converged stress function is obtained based on a linear combination of power functions fitted by the software.

[0015] Optionally, the converged stress function is as follows:

[0016] σ x =c0+c1x -1 +c2x -2 +……+c m x -m ;

[0017] Wherein, m=n-1.

[0018] Optionally, the limit value of the convergence stress function is used as the convergence stress of the aircraft structure, including:

[0019] When x takes the convergence extreme value in the convergence region, σ x This is the convergence stress of the aircraft structure.

[0020] Optionally, the element type in the finite element model is a second-order tetrahedral element.

[0021] A second aspect of the present invention provides a device for acquiring convergence stress of an aircraft structure body, for executing the method as described in any one of the first aspects.

[0022] A third aspect of the present invention provides a computer storage medium, comprising: a memory and a processor;

[0023] The memory is configured to store executable instructions;

[0024] The processor is configured to implement the method as described in any one of the first aspects when executing the executable instructions stored in the memory.

[0025] The present invention provides a method and device for obtaining the convergent stress of an aircraft structure body. The method provided by the present invention can accurately and quickly calculate the convergent stress of the aircraft structure body, meeting the strength calculation and verification requirements of the aircraft structure body. Compared with traditional design methods, the present invention improves the calculation accuracy. The method provided by the present invention is simple to operate, realizes the standardization of calculation methods and processes, avoids large errors in results when different researchers perform calculations, and provides aircraft designers with a method for accurately calculating the convergent stress of an aircraft structure body. DETAILED DESCRIPTION

[0026] To make the objectives, technical solutions, and advantages of the embodiments of the present invention more clear, the technical solutions in the embodiments of the present invention are clearly and completely described below. Obviously, the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making any creative efforts are within the scope of protection of the present invention.

[0027] The features and illustrative embodiments of various aspects of the present invention will be described in detail below. In the detailed description below, many specific details are proposed in order to provide a comprehensive understanding of the present invention. However, it will be apparent to those skilled in the art that the present invention can be implemented without the need for some of these specific details. The following description of the embodiments is merely intended to provide a better understanding of the present invention by illustrating examples of the present invention. The present invention is in no way limited to any specific arrangement and method proposed below, but rather encompasses any improvements, replacements, and modifications to structures, methods, and devices without departing from the spirit of the present invention. In the following description, known structures and techniques are not shown to avoid unnecessary ambiguity in the present invention.

[0028] In the description of the present invention, it should be noted that the use of terms such as "center," "upper," "lower," "left," "right," "vertical," "horizontal," "inner," and "outer" to indicate directions or positional relationships is intended solely to facilitate and simplify the description of the present invention and should not be construed as limiting the present invention. Furthermore, the use of ordinal numbers (e.g., "first and second," etc.) is intended to distinguish between objects and is not intended to limit the order of the objects. Such terms should not be construed as indicating or implying relative importance.

[0029] In the description of the present invention, it should be noted that, unless otherwise specified or limited, the terms "mounted," "connected," and "connected" should be understood broadly, and may refer to direct connection or indirect connection through an intermediary. Those skilled in the art will understand the specific meanings of these terms in the present invention.

[0030] It should be noted that, in the absence of conflict, the embodiments of the present invention and the features therein may be combined with each other, and the various embodiments may refer to and quote each other.

[0031] The present invention will be further described in detail below with reference to examples, but the embodiments of the present invention are not limited thereto.

[0032] The present invention performs static strength calculation based on the finite element analysis method. First, the minimum thickness t of the structure body is used as the reference size and the unit type is the second-order tetrahedron unit to calculate the maximum comprehensive stress of the structure body, which is recorded as σ1. Secondly, 1 / 2, 1 / 3, 1 / 4...1 / n (n≥2) of the minimum thickness t of the structure body is used as the reference size and the unit type is the second-order tetrahedron unit to calculate the maximum comprehensive stress of the structure body, which is recorded as σ2, σ3, σ4...σ n (n≥2), when (σ n -σ n-1 ) / σ n-1 When ≤0.02, the calculation is terminated; finally, according to σ1 to σ n The software is used to fit the convergence function based on the n calculation data. The limit value of the function is the convergence stress of the structure. The present invention can conveniently, quickly and accurately calculate the convergence stress of the structure.

[0033] The present invention provides a method for calculating convergence stress of an aircraft structure, which mainly comprises the following steps:

[0034] S1. Taking the minimum thickness t of the structure as the reference dimension, establish a finite element model of the aircraft structure body, using the second-order tetrahedron element as the unit type;

[0035] S2. Based on the boundary conditions and load conditions, the maximum stress calculation result of the aircraft structure body is obtained, which is recorded as σ1;

[0036] S3. Using 1 / 2, 1 / 3, 1 / 4, ..., 1 / n (n ≥ 2) of the minimum thickness t of the aircraft structure as reference dimensions, establish a finite element model of the aircraft structure, using a second-order tetrahedron element as the unit type;

[0037] S4. According to the boundary conditions and load conditions, the maximum stress calculation results of the aircraft structure body with 1 / 2, 1 / 3, 1 / 4...1 / n (n≥2) of the minimum thickness t as the reference size are recorded as σ2, σ3, σ4...σ n(n≥2), when (σ n -σ n-1 ) / σ n-1 When ≤0.02, terminate the calculation;

[0038] S5, according to σ1 to σ n The n calculation data are (1, σ1), (2, σ2) ... (n, σ n ) are the horizontal and vertical coordinates, and the linear combination of power functions is fitted by software. The linear combination of power functions can be expressed as:

[0039] σ x =c0+c1x -1 +c2x -2 +……+c m x -m

[0040] m=n-1, obtained according to the software fitting results;

[0041] S6. When x takes the convergence extreme value in the convergence domain, σ x This is the convergence stress of the aircraft structure.

[0042] It can be understood that when the final value of n is 2, the convergence stress function is convergence stress = σ2.

[0043] The present invention provides a method for obtaining the convergent stress of an aircraft structure body, which performs calculation and analysis based on the finite element method, with the aim of conveniently, quickly and accurately calculating the stress of the aircraft structure body.

[0044] Taking the rear flange of an aircraft refueling device as an example, the rear flange is made of 7050 aluminum alloy with a minimum thickness of 6 mm. The calculation method of the convergent stress of the rear flange includes the following steps:

[0045] S1. The minimum thickness of the rear butt flange is 6mm as the reference size, and the finite element model of the rear butt flange is established using finite element analysis software;

[0046] S2. According to the boundary conditions and load conditions, the maximum stress calculation result of the rear butt flange σ1 is 220.80MPa;

[0047] S3. Taking 3mm as the reference size, the finite element model of the rear butt flange was established. According to the boundary conditions and load conditions, the maximum stress calculation result σ2 of the rear butt flange with 3mm as the reference size was obtained to be 306.72MPa.

[0048] If S4 and (σ2-σ1) / σ1>0.02 do not meet the requirements, the finite element model of the rear butt flange is established with 2mm as the reference size. According to the boundary conditions and load conditions, the maximum stress calculation result σ3 of the rear butt flange with 2mm as the reference size is 329.19MPa.

[0049] If S5 and (σ3-σ2) / σ2>0.02 do not meet the requirements, the finite element model of the rear butt flange is established with 1.5mm as the reference size. According to the boundary conditions and load conditions, the maximum stress calculation result σ4 of the rear butt flange with 1.5mm as the reference size is 339.54MPa.

[0050] S6, (σ4-σ3) / σ3>0.02, which does not meet the requirements, then 1.2mm is used as the reference size to establish the rear butt flange finite element model. According to the boundary conditions and load conditions, the maximum stress calculation result σ5 of the rear butt flange with 1.2mm as the reference size is 345.50MPa;

[0051] S7, (σ5-σ4) / σ4≤0.02, meets the requirements. According to the five calculated data of σ1 to σ5, with (1, 220.80), (2, 306.72), (3, 329.19), (4, 339.54), and (5, 345.50) as coordinates, the convergent stress function is fitted:

[0052] σ x =367.72-105.26x -1 -26.56x -2 -12.52x -3 -2.58x -4

[0053] S8. The limit value of the convergence stress function is 367.72 MPa, that is, the convergence stress of the rear butt flange is 367.72 MPa.

[0054] The above specific implementation methods are detailed descriptions of the present invention. It cannot be considered that the specific implementation methods of the present invention are limited to these descriptions. For ordinary technicians in the technical field to which the present invention belongs, they can make several simple deductions and substitutions without departing from the concept of the present invention, which should be regarded as falling within the scope of protection of the present invention.

Claims

1. A method for obtaining convergence stress of an aircraft structure, characterized in that: The main steps include: S1. Using the minimum thickness t of the aircraft structure as a reference dimension, a finite element model of the aircraft structure is established using finite element analysis software; S2. According to the boundary conditions and load conditions, the maximum stress calculation result of the aircraft structure body is obtained, which is recorded as σ1; let the value of n be 2; S3. Taking 1 / n of the minimum thickness t of the aircraft structure as the reference size, the finite element model of the aircraft structure is established using finite element analysis software. According to the boundary conditions and load conditions, the maximum stress calculation result of the aircraft structure with 1 / n of the minimum thickness t as the reference size is obtained and recorded as σ n ; S4, judge whether it satisfies (σ n -σ n-1 ) / σ n-1 ≤0.

02. If not satisfied, add 1 to the value of n and return to S3. If satisfied, proceed to S5. S5, according to σ1 to σ n The convergent stress function is fitted based on n calculation data, and the limit value of the convergent stress function is used as the convergent stress of the aircraft structure.

2. The method for obtaining convergence stress of an aircraft structure according to claim 1, characterized in that: When the final value of n is 2, the convergence stress function is convergence stress = σ n .

3. The method for obtaining convergence stress of an aircraft structure according to claim 1, characterized in that: When the final value of n is greater than 2, the converged stress function is obtained based on the linear combination of power functions fitted by the software.

4. The method for obtaining convergence stress of an aircraft structure according to claim 3, characterized in that: The converged stress function is as follows: σ x =c0+c1x -1 +c2x -2 +……+c m x -m ; Wherein, m=n-1.

5. The method for obtaining convergence stress of an aircraft structure according to claim 4, characterized in that: The limit value of the convergent stress function is used as the convergent stress of the aircraft structure, including: When x takes the convergence extreme value in the convergence region, σ x This is the convergence stress of the aircraft structure.

6. The method for obtaining convergence stress of an aircraft structure according to claim 1, characterized in that: The element type in the finite element model is second-order tetrahedral element.

7. A device for obtaining convergence stress of an aircraft structure, characterized in that: Used to perform the method according to any one of claims 1 to 6.

8. A computer storage medium, characterized in that include: memory and processor; The memory is configured to store executable instructions; The processor is configured to implement the method according to any one of claims 1 to 6 when executing the executable instructions stored in the memory.