Metal material stress plasticity correction method
By employing linear finite element analysis and a modified Ramberg-Osgood nonlinear hardening exponent, the stress concentration problem in aircraft structures was solved, achieving efficient and accurate plasticity correction of metallic materials and improving analysis efficiency and accuracy.
Patent Information
- Application Number
- CN202511721289.0
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2025-11-21
- Publication Date
- 2026-03-03
AI Technical Summary
In existing technologies for aircraft structural strength analysis, finite element analysis requires significant computational resources and is time-consuming, while engineering analysis methods are not very accurate when stress exceeds the material's strength limit and are difficult to effectively correct for the plastic strain of metallic materials.
The stress distribution was obtained by linear finite element static strength analysis. By modifying the Ramberg-Osgood nonlinear hardening index and combining material parameters such as yield strength, tensile strength and elongation, a stress equation was constructed for plastic correction to obtain the true stress.
It improves the efficiency and accuracy of plastic correction at stress concentration points in aircraft metal structures, meets design requirements, and reduces computational resources and time requirements.
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Figure CN121598684A_ABST
Abstract
Description
Technical Field
[0001] This application belongs to the field of aircraft structural strength analysis technology, and specifically relates to a method for stress-plasticity correction of metallic materials. Background Technology
[0002] In aircraft structural strength analysis, the structure is required to not yield under restrained loads and not fail under ultimate loads. Typically, the safety factor for ultimate loads is 1.5. For metal structures such as landing gear, due to unavoidable stress concentration, the stress at localized stress concentration points may exceed the material's strength limit under ultimate loads.
[0003] However, because commonly used materials in landing gear structures, such as aluminum alloys and high-strength steel, have good plasticity, structural failure is rare. In such cases, plastic correction can be applied to the structural stress. Common methods include finite element analysis considering material nonlinearity or engineering methods. However, due to the large model size, finite element analysis requires significant computational resources and time. Engineering methods can use the Neuberger criterion and the Ramberg-Osgood nonlinear stress-strain curve for plastic correction. However, the Ramberg-Osgood nonlinear hardening exponent n from the material handbook used in engineering methods is only relatively accurate for calculating plastic strain near the yield strength; its accuracy is low when the stress exceeds the material's strength limit.
[0004] Therefore, there is an urgent need for a technical solution to overcome or mitigate at least one of the aforementioned defects in the existing technology. Summary of the Invention
[0005] The purpose of this application is to provide a method for correcting the stress-plasticity of metallic materials to solve at least one problem existing in the prior art.
[0006] The technical solution of this application is:
[0007] The first aspect of this application provides a method for stress-plasticity correction of metallic materials, comprising:
[0008] Step 1: Obtain the material parameters of the aircraft's metallic structure;
[0009] Step 2: Using the linear finite element static strength analysis method, obtain the stress distribution of the aircraft metal material structure under ultimate load. If the stress meets the preset conditions, proceed to Step 3.
[0010] Step 3: Correct the Ramberg-Osgood nonlinear hardening index at tensile failure based on the material parameters to obtain the corrected nonlinear hardening index;
[0011] Step 4: Construct the stress equation and substitute the nonlinear hardening index into the stress equation to obtain the true stress of the aircraft metal material structure after plastic correction.
[0012] In a preferred embodiment of this application, in step one, the material parameters include: yield strength σ. 0.2 Tensile limit σ b Young's modulus E, elongation δ.
[0013] In a preferred embodiment of this application, in step two, the preset condition is the maximum stress σ at the stress concentration point. max Exceeding the strength limit of metallic materials.
[0014] In a preferred embodiment of this application, in step three, based on the yield strength σ... 0.2 Tensile limit σ b The elongation δ is used to correct the Ramberg-Osgood nonlinear hardening index at tensile failure, resulting in the corrected nonlinear hardening index:
[0015] ;
[0016] in, This is the corrected nonlinear hardening exponent.
[0017] In a preferred embodiment of this application, in step four, the stress equation is:
[0018] ;
[0019] Where σ is the actual stress.
[0020] A second aspect of this application provides a stress-plasticity correction system for metallic materials, comprising:
[0021] The material parameter acquisition module is used to acquire the material parameters of the aircraft's metallic material structure.
[0022] The static strength analysis module is used to obtain the stress distribution of the aircraft metal material structure under ultimate load using the linear finite element static strength analysis method. If the stress meets the preset conditions, it will enter the nonlinear hardening index correction module.
[0023] The nonlinear hardening index correction module is used to correct the Ramberg-Osgood nonlinear hardening index at tensile failure based on the material parameters, so as to obtain the corrected nonlinear hardening index.
[0024] The real stress calculation module is used to construct the stress equation. The nonlinear hardening index is substituted into the stress equation to obtain the real stress of the aircraft metal material structure after plastic correction.
[0025] In a preferred embodiment of this application, the material parameter acquisition module includes the following material parameters: yield strength σ. 0.2 Tensile limit σ b Young's modulus E, elongation δ.
[0026] In a preferred embodiment of this application, in the static strength analysis module, the preset condition is the maximum stress σ at the stress concentration point. max Exceeding the strength limit of metallic materials.
[0027] In a preferred embodiment of this application, the nonlinear hardening exponent correction module is based on the yield strength σ. 0.2 Tensile limit σ b The elongation δ is used to correct the Ramberg-Osgood nonlinear hardening index at tensile failure, resulting in the corrected nonlinear hardening index:
[0028] ;
[0029] in, This is the corrected nonlinear hardening exponent.
[0030] In a preferred embodiment of this application, the stress equation in the real stress calculation module is:
[0031] ;
[0032] Where σ is the actual stress.
[0033] The invention has at least the following beneficial technical effects:
[0034] The stress-plastic correction method for metallic materials proposed in this application can conveniently and quickly perform plastic correction on the stress of metallic materials in aircraft structures, solving the problems of low calculation efficiency and low accuracy of conventional finite element methods when the stress at stress concentration points exceeds the material's strength limit. Attached Figure Description
[0035] Figure 1 This is a flowchart of a method for stress-plasticity correction of metallic materials according to one embodiment of this application. Detailed Implementation
[0036] To make the objectives, technical solutions, and advantages of this application clearer, the technical solutions in the embodiments of this application will be described in more detail below with reference to the accompanying drawings. In the drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments are some, but not all, embodiments of this application. The embodiments described below with reference to the accompanying drawings are exemplary and intended to explain this application, and should not be construed as limiting this application. All other embodiments obtained by those skilled in the art based on the embodiments of this application without creative effort are within the scope of protection of this application. The embodiments of this application will be described in detail below with reference to the accompanying drawings.
[0037] In the description of this application, it should be understood that the terms "center", "longitudinal", "lateral", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limiting the scope of protection of this application.
[0038] The following is in conjunction with the appendix Figure 1 This application will be described in further detail.
[0039] The first aspect of this application provides a method for correcting the stress-plasticity of metallic materials, comprising the following steps:
[0040] Step 1: Obtain the material parameters of the aircraft's metallic structure;
[0041] Step 2: Use the linear finite element static strength analysis method to obtain the stress distribution of the aircraft metal material structure under the ultimate load. If the stress meets the preset conditions, proceed to step 3.
[0042] Step 3: Correct the Ramberg-Osgood nonlinear hardening index at tensile failure based on the material parameters to obtain the corrected nonlinear hardening index;
[0043] Step 4: Construct the stress equation and substitute the nonlinear hardening exponent into the stress equation to obtain the true stress of the aircraft metal material structure after plastic correction.
[0044] The stress-plasticity correction method for metallic materials in this application firstly determines the materials used in the aircraft's metallic structure in step one. In the strength analysis of the aircraft's metallic structure, for ductile metallic materials, material parameters, including the yield strength σ, are obtained from a material handbook. 0.2 Tensile limit σb Young's modulus E, elongation δ.
[0045] Then, in step two, the conventional linear finite element static strength analysis method is used to obtain the stress distribution of the aircraft's metallic material structure under ultimate load. If the maximum stress σ at the stress concentration point is... max If the strength limit of the metallic material is exceeded, the subsequent plasticity correction step is initiated.
[0046] In step three, based on the yield strength σ 0.2 Tensile limit σ b The elongation δ is used to correct the Ramberg-Osgood nonlinear hardening index of aircraft metallic structures at tensile failure, resulting in the corrected nonlinear hardening index:
[0047] ;
[0048] in, This is the corrected nonlinear hardening exponent.
[0049] Finally, in step four, the following stress equations are solved simultaneously using an iterative method:
[0050] ;
[0051] Where σ is the actual stress.
[0052] Substituting the nonlinear hardening exponent into the stress equation yields the true stress of the aircraft metal material structure after plastic correction.
[0053] In one specific embodiment of this application, the steps for performing static strength analysis on an aircraft structure to achieve plastic correction are as follows:
[0054] The aircraft's metallic structure is made of 0Cr13Ni8Mo2Al material. The yield strength of this material was obtained from the material handbook. Tensile limit Young's modulus E=192000MPa, elongation δ=10%.
[0055] Linear finite element static strength analysis was performed on the metallic structure of the aircraft to obtain the maximum stress at the stress concentration point. The maximum stress has exceeded the tensile limit of the material. The structural stress should be plastically corrected according to the following method.
[0056] According to the yield strength σ 0.2 Tensile limit σ b The elongation δ is used to correct the Ramberg-Osgood nonlinear hardening index of the material at tensile failure, resulting in the corrected nonlinear hardening index. :
[0057] ;
[0058] Solve the following stress equations simultaneously using an iterative method:
[0059] ;
[0060] The true stress of the structure after plastic correction is The strength meets the design requirements.
[0061] The stress-plasticity correction method for metallic materials proposed in this application considers the elongation and ultimate tensile strength of the material. After correcting the Ramberg-Osgood nonlinear hardening exponent n of the material with the elongation and ultimate tensile strength, the method is substituted into the Neuberger criterion to calculate the corresponding plastic stress and strain. This method can conveniently, quickly, and accurately perform plastic correction on high stresses exceeding the material's ultimate tensile strength at stress concentration points in metallic structures, thereby improving analysis efficiency and accuracy.
[0062] Based on the above-described method for correcting stress-plasticity in metallic materials, a second aspect of this application provides a system for correcting stress-plasticity in metallic materials, comprising:
[0063] The material parameter acquisition module is used to acquire the material parameters of the aircraft's metallic material structure.
[0064] The static strength analysis module is used to obtain the stress distribution of aircraft metal material structures under ultimate load using the linear finite element static strength analysis method. If the stress meets the preset conditions, it will enter the nonlinear hardening index correction module.
[0065] The nonlinear hardening index correction module is used to correct the Ramberg-Osgood nonlinear hardening index at tensile failure based on material parameters, and obtain the corrected nonlinear hardening index.
[0066] The Real Stress Calculation Module is used to construct stress equations and substitute the nonlinear hardening exponent into the stress equations to obtain the real stress of the aircraft metal material structure after plastic correction.
[0067] In a preferred embodiment of this application, the aircraft metal material structure can be a 0Cr13Ni8Mo2Al component.
[0068] In a preferred embodiment of this application, the material parameter acquisition module includes the following material parameters: yield strength σ. 0.2 Tensile limit σ b Young's modulus E, elongation δ.
[0069] In a preferred embodiment of this application, the static strength analysis module is preset with the maximum stress σ at the stress concentration point as the predefined condition. max Exceeding the strength limit of metallic materials.
[0070] In a preferred embodiment of this application, the nonlinear hardening exponent correction module is based on the yield strength σ. 0.2 Tensile limit σ b The elongation δ is used to correct the Ramberg-Osgood nonlinear hardening index at tensile failure, resulting in the corrected nonlinear hardening index:
[0071] ;
[0072] in, This is the corrected nonlinear hardening exponent.
[0073] In a preferred embodiment of this application, the stress equation in the real stress calculation module is:
[0074] ;
[0075] Where σ is the actual stress.
[0076] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.
Claims
1. A method for correcting stress-plasticity in metallic materials, characterized in that, include: Step 1: Obtain the material parameters of the aircraft's metallic structure; Step 2: Using the linear finite element static strength analysis method, obtain the stress distribution of the aircraft metal material structure under ultimate load. If the stress meets the preset conditions, proceed to Step 3. Step 3: Correct the Ramberg-Osgood nonlinear hardening index at tensile failure based on the material parameters to obtain the corrected nonlinear hardening index; Step 4: Construct the stress equation and substitute the nonlinear hardening index into the stress equation to obtain the true stress of the aircraft metal material structure after plastic correction.
2. The method for correcting stress and plasticity in metallic materials according to claim 1, characterized in that, In step one, the material parameters include: yield strength σ 0.2 Tensile limit σ b Young's modulus E, elongation δ.
3. The method for correcting stress and plasticity in metallic materials according to claim 2, characterized in that, In step two, the preset condition is the maximum stress σ at the stress concentration point. max Exceeding the strength limit of metallic materials.
4. The method for correcting stress and plasticity in metallic materials according to claim 3, characterized in that, In step three, based on the yield strength σ 0.2 Tensile limit σ b The elongation δ is used to correct the Ramberg-Osgood nonlinear hardening index at tensile failure, resulting in the corrected nonlinear hardening index: ； in, This is the corrected nonlinear hardening exponent.
5. The method for correcting stress and plasticity in metallic materials according to claim 4, characterized in that, In step four, the stress equation is: ； Where σ is the actual stress.
6. A stress-plasticity correction system for metallic materials, characterized in that, include: The material parameter acquisition module is used to acquire the material parameters of the aircraft's metallic material structure. The static strength analysis module is used to obtain the stress distribution of the aircraft metal material structure under ultimate load using the linear finite element static strength analysis method. If the stress meets the preset conditions, it will enter the nonlinear hardening index correction module. The nonlinear hardening index correction module is used to correct the Ramberg-Osgood nonlinear hardening index at tensile failure based on the material parameters, so as to obtain the corrected nonlinear hardening index. The real stress calculation module is used to construct the stress equation. The nonlinear hardening index is substituted into the stress equation to obtain the real stress of the aircraft metal material structure after plastic correction.
7. The stress-plasticity correction system for metallic materials according to claim 6, characterized in that, In the material parameter acquisition module, the material parameters include: yield strength σ 0.2 Tensile limit σ b Young's modulus E, elongation δ.
8. The stress-plasticity correction system for metallic materials according to claim 7, characterized in that, In the static strength analysis module, the preset condition is the maximum stress σ at the stress concentration point. max Exceeding the strength limit of metallic materials.
9. The stress-plasticity correction system for metallic materials according to claim 8, characterized in that, In the nonlinear hardening exponent correction module, based on the yield strength σ 0.2 Tensile limit σ b The elongation δ is used to correct the Ramberg-Osgood nonlinear hardening index at tensile failure, resulting in the corrected nonlinear hardening index: ； in, This is the corrected nonlinear hardening exponent.
10. The stress-plasticity correction system for metallic materials according to claim 9, characterized in that, In the actual stress calculation module, the stress equation is: ； Where σ is the actual stress.