A method for optimizing the ply proportion of a composite laminate

Abstract

The application belongs to the field of composite structure layer design of aircraft structure, and particularly relates to a composite laminate layer proportion optimization method, step S1: in the aircraft design process, the design load and the aircraft structure layout are determined, the design allowable value of the composite material applied to the aircraft is determined, including the axial allowable strain epsilon ad , the shear allowable strain gamma ad ; step S2: based on the external load of the aircraft structure and the specific layout, the load characteristic value T of the aircraft to be layered structure is calculated, and the load characteristic value T represents the trend characteristics of the axial load and the shear load of the aircraft to be layered structure; step S3: a plurality of preset proportion composite laminates in a layering library are obtained, and the material characteristic value T m of the composite laminates is calculated; step S4: based on the structure load characteristic value T, the composite laminate with the material characteristic value T m matching the load characteristic value T is selected. The relationship between the structure and the material is quantified, so that the structure material selection is more intuitive and optimal.

Description

Technical Field

[0001] This application belongs to the field of composite structure layup design for aircraft structures, and specifically relates to a method for optimizing the layup ratio of composite laminates. Background Technology

[0002] In existing technologies, the optimization of ply ratios has failed to conduct precise quantitative analysis of structural load-bearing characteristics, failed to define a mathematical model between structural load-bearing capacity and ply ratio, failed to establish a clear relationship between the two, and failed to enable the structure to accurately match a better ply ratio.

[0003] In composite material structure design, the traditional design concept is to determine the ply ratio based on experience in bearing tensile, compressive, and shear loads within the structural plane. However, experience-based design is a general design principle. Aircraft structures exhibit different structural load characteristics due to variations in shape, load, and layout. For example, the load distribution among the main load-bearing components of a wing—whether it's a beam wing, a single-piece wing, or a multi-wall wing—can differ significantly.

[0004] To address the unique load-bearing characteristics resulting from the different layouts of each aircraft type, targeted ply design is proposed. The concept of "load-bearing characteristic value" is used to define the structural load-bearing characteristics and to quantitatively calculate the tensile, compressive, and shear loads within the structural plane. "E and G value design of composite laminates based on structural characteristics" is proposed, which involves optimizing the correlation between load-bearing characteristic value and laminate ply ratio, matching the optimal ply, and dividing the structure into regions according to ply ratio based on load-bearing characteristic value. Summary of the Invention

[0005] To address the aforementioned problems, a method for optimizing the layup ratio of composite laminates is provided. This method is applied to the ply structure of an aircraft, aiming to perform targeted composite material layup based on load-bearing capacity. Traditional methods rely on experience to lay up the composite material at different angles and in varying proportions to create a good, lightweight structure. The composite laminate layup ratio optimization method includes:

[0006] A method for optimizing the layup ratio of composite laminates includes:

[0007] Step S1: During the aircraft design process, determine the design loads and aircraft structural layout, and clarify the allowable design values ​​for composite materials used in the aircraft, including the allowable axial strain ε. ad Allowable shear strain γ ad ;

[0008] Step S2: Based on the external loads and specific layout of the aircraft structure, calculate the load-bearing characteristic value T of the aircraft structure to be laid, wherein the load-bearing characteristic value T characterizes the trend characteristics of the axial load and shear load of the aircraft structure to be laid;

[0009] Step S3: Obtain multiple composite laminates with preset proportions from the layup library, and calculate the material characteristic value T of the composite laminates. m ;

[0010] Step S4: Based on the structural bearing capacity characteristic value T, select a material characteristic value T that matches the bearing capacity characteristic value T. m Composite material laminate.

[0011] Preferably, the method for calculating the bearing characteristic value T is as follows:

[0012] Step S21: Preset the initial elastic modulus E x Initial shear modulus G xy ;

[0013] Step S22: Based on the initial elastic modulus E x Initial shear modulus G xy Obtain the corresponding initial axial strain ε x , with shear strain γ xy ;

[0014] Step S23: Iterate through the elastic modulus E x Shear modulus G xy , causing axial strain ε x With shear strain γ xy The ratio |ε x / γ xy |Approximately allowable strain ε ad With allowable stress γ ad The ratio ε ad / γ ad Under the preset conditions, obtain the elastic modulus E after iteration. x n With shear modulus G xy n ;

[0015] Step S24: Based on the iterative elastic modulus E x n With shear modulus G xy n Calculate the bearing characteristic value T.

[0016] Preferably, the axial strain ε x Shear strain γ xyIncluding multiple, specifically: the aircraft to be laid layer structure is divided into multiple units, the load-bearing characteristic value T of each unit is calculated, and a matching composite material laminate is selected based on the load-bearing characteristic value T of each unit.

[0017] Preferably, the specific formula for calculating the bearing characteristic value T is as follows: When ε x n If the strain is tensile, then T = |T| is the tensile-shear characteristic value; if ε x n If the strain is compressive, then T = -|T| is the characteristic value of compressive shear.

[0018] Among them, the axial strain ε after iteration x n The shear strain γ after iteration xy n The axial stress σ after iteration x n The Poisson ratio μ after iteration yx n The normal stress σ after iteration y n The iterative shear stress τ xy n ;

[0019] The specific process of calculating T is as follows:

[0020] Step 1: First, based on the plane stress state relationship of orthotropic materials:

[0021]

[0022] in,

[0023]

[0024]

[0025]

[0026] Conclusion:

[0027] |ε x / γ xy | represents the ratio of axial strain to shear strain. In this paper, T is defined as the element bearing characteristic value, which represents a measure of the magnitude and proportion of the axial load and shear load of the structure. If ε x If the strain is tensile, then T = |T| is the tensile-shear characteristic value; if ε xIf the strain is compressive, then T = -|T| is the characteristic value of compression-shear; the smaller |T| is, the more the structural element is mainly subjected to shear; the larger |T| is, the more the structural element is mainly subjected to tension or compression.

[0028] Step 2: Select an existing plywood library of composite laminates for E x G xy The value is given, and the allowable strain ε is also given. ad γ ad And through the E of composite laminate x G xy The value, ε, is obtained through finite element analysis and calculation, corresponding to the axial strain. x Shear strain γ xy axial stress σ x Poisson's ratio μ yx Normal stress σ y Shear stress τ xy ;

[0029] Formula 1 can be written as:

[0030]

[0031] Other:

[0032]

[0033] Right now:

[0034]

[0035] E x ′ is E x The first iteration value, G xy ′ is G xy The first iteration value;

[0036] Substituting formula 2 into formula 1 for one iteration, where,

[0037]

[0038] σ x ′,μ yx ′,σ y ′,τ xy ′,ε x ′, γ xy ′ are all in E x ′ and G xy Based on the first iteration value obtained through finite element analysis or calculation, the value is obtained from the first iteration.

[0039] If not satisfied The set conditions are then used for a second iteration.

[0040]

[0041] According to Formula 2, also:

[0042]

[0043] Right now:

[0044]

[0045] Substituting Formula 3 into Formula 1 for a second iteration follows the same principle as the first iteration.

[0046]

[0047] This process can be repeated multiple times until the working strain ratio |ε| is reached. y n / γ xy n | The ratio ε that is infinitely close to the design allowable value ad / γ ad In this way, the stiffness trimming of the composite material perfectly matches the structural load-bearing characteristics, resulting in relatively high material design efficiency. Therefore, this is the process of matching the mechanical properties of the material with the load-bearing characteristics of the structure. The final E... x n G xy n The optimal ratio is given by the superscripts, which represent the number of iterations. The calculation principle is to assume that a composite laminate with a certain layup ratio is regarded as a specific single-layer composite material, and to design a composite material with matching performance based on the structural load-bearing characteristics, that is, to design the E and G values ​​of the composite laminate.

[0048] Preferably, the characteristic values ​​T of different composite laminates in a preset optional layup library are used. m Calculate the applicable boundary characteristic value T for each composite laminate. b The material characteristic value T of composite laminate m The characteristic value T of the composite laminate boundary represents the optimal ratio of axial and shear bearing capacity for this specific proportion of the laminate. b This characterizes the applicable range of axial and shear bearing capacity of this specific ratio of laminate. When the structural bearing characteristic value T is within a certain range of the laminate boundary characteristic value T... b Within the corresponding range, select the boundary feature value T. b The corresponding material characteristic value T m and the corresponding composite laminate materials;

[0049] T m The calculation formula is: Where E 1材G represents the axial modulus of the composite laminate material. 1材 The shear modulus of the composite laminate material; material characteristic value T b =kT m Where k is a correction factor. The principle is that in structural design, due to various reasons (including process constraints, design experience, and experimental accumulation), designers often choose commonly used ply ratios from the ply library, but they are unsure how to select a ply ratio from the commonly used ply library that matches the load-bearing characteristics. To solve this problem, laminates with specific ply ratios are considered as anisotropic materials. Material characteristic values ​​for the laminate are defined, and boundary critical values, i.e., boundary characteristic values ​​T, are established. b A correction factor is introduced. The zoning critical value divides the laminate of each ply ratio into an interval in which the load-bearing characteristic value |T| applies.

[0050] Preferably, the initial axial strain ε x , with shear strain γ xy It was obtained through finite element analysis.

[0051] Preferably, the load-bearing characteristic value T and the material characteristic value T m With boundary eigenvalue T b Range matching, specifically, involves selecting the material feature value T that is closest to |T|. m And less than the material's characteristic value T m The corresponding boundary eigenvalue T b .

[0052] The advantages of this application include: establishing the relationship between the mechanical requirements and materials of the aircraft ply structure by proposing load-bearing characteristic values ​​and material characteristic values; and making the plying of the aircraft ply structure more targeted through digital quantitative characterization. When plying the aircraft ply structure, the corresponding plying method can be selected according to the different loads of different parts of the aircraft ply structure; and realizing the correlation optimization between load-bearing characteristic values ​​and the ply ratio of laminates to match the optimal ply. Attached Figure Description

[0053] Figure 1 This is a flowchart of a preferred embodiment of the composite material laminate layup ratio optimization method of this application. Detailed Implementation

[0054] 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 only some, not all, of the 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 inventive 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.

[0055] To address the aforementioned problems, a method for optimizing the layup ratio of composite laminates is provided. This method is applied to the ply structure of an aircraft, aiming to perform targeted composite material layup based on load-bearing capacity. Traditional methods rely on experience to lay up the composite material at different angles and in varying proportions to create a good, lightweight structure. The composite laminate layup ratio optimization method includes:

[0056] A method for optimizing the layup ratio of composite laminates, characterized by comprising:

[0057] Step S1: During the aircraft design process, determine the design loads and aircraft structural layout, and clarify the allowable design values ​​for composite materials used in the aircraft, including the allowable axial strain ε. ad Allowable shear strain γ ad ;

[0058] Step S2: Based on the external loads and specific layout of the aircraft structure, calculate the load-bearing characteristic value T of the aircraft structure to be laid, wherein the load-bearing characteristic value T characterizes the trend characteristics of the axial load and shear load of the aircraft structure to be laid;

[0059] Step S3: Obtain multiple composite laminates with preset proportions from the layup library, and calculate the material characteristic value T of the composite laminates. m ;

[0060] Step S4: Based on the structural bearing capacity characteristic value T, select a material characteristic value T that matches the bearing capacity characteristic value T. m Composite material laminate.

[0061] In some optional implementations, the method for calculating the bearing characteristic value T is as follows:

[0062] Step S21: Preset the initial elastic modulus E x Initial shear modulus G xy ;

[0063] Step S22: Based on the initial elastic modulus E x Initial shear modulus G xy Obtain the corresponding initial axial strain ε x , with shear strain γ xy ;

[0064] Step S23: Iterate through the elastic modulus E x Shear modulus G xy , causing axial strain ε x With shear strain γ xy The ratio |ε x / ε xy |Approximately allowable strain ε ad With allowable stress γ ad The ratio ε ad / γ ad Under the preset conditions, obtain the elastic modulus E after iteration. x n With shear modulus G xy n .

[0065] Step S24: Based on the iterative elastic modulus E x n With shear modulus G xy n Calculate the bearing characteristic value T.

[0066] In some alternative embodiments, the axial strain ε x Shear strain γ xy Including multiple types, specifically: axial strain ε x With shear strain γ xy The process involves dividing the aircraft's structure to be laid into multiple units, calculating the load-bearing characteristic value T for each unit, and selecting matching composite material laminates based on the load-bearing characteristic value T for each unit.

[0067] In some optional implementations, the specific formula for calculating the bearing characteristic value T is as follows: When ε x n If the strain is tensile, then T = |T| is the tensile-shear characteristic value; if ε x n If the strain is compressive, then T = -|T| is the characteristic value of compressive shear.

[0068] Among them, the axial strain ε after iteration x n The shear strain γ after iteration xy n The axial stress σ after iteration xn The Poisson ratio μ after iteration yx n The normal stress σ after iteration y n The iterative shear stress τ xy n ;

[0069] The specific process of calculating T is as follows:

[0070] Step 1: First, based on the plane stress state relationship of orthotropic materials:

[0071]

[0072] in,

[0073]

[0074]

[0075]

[0076] Conclusion:

[0077] |ε x / γ xy | represents the ratio of axial strain to shear strain. In this paper, T is defined as the element bearing characteristic value, which represents a measure of the magnitude and proportion of the axial load and shear load of the structure. If ε x If the strain is tensile, then T = |T| is the tensile-shear characteristic value; if ε x For compressive strain, T = -|T| is the characteristic value of compression-shear. The smaller |T| is, the more the structural element is mainly subjected to shear; the larger |T| is, the more the structural element is mainly subjected to tension or compression.

[0078] Step 2: Select an existing plywood library of composite laminates for E x G xy The value is given, and the allowable strain ε is also given. ad γ ad And through the E of composite laminate x G xy The value, ε, is obtained through finite element analysis and calculation, corresponding to the axial strain. x Shear strain γ xy axial stress σ x Poisson's ratio μ yx Normal stress σ y Shear stress τ xy ;

[0079] Formula 1 can be written as:

[0080]

[0081] Other:

[0082]

[0083] Right now:

[0084]

[0085] E x ′ is E x The first iteration value, G xy ′ is G xy The first iteration value;

[0086] Substituting formula 2 into formula 1 for one iteration, where,

[0087]

[0088] σ x ′,μ yx ′,σ y ′,τ xy ′,ε x ′, γ xy ′ are all in E x ′ and G xy Based on the first iteration value obtained through finite element analysis or calculation, the value is obtained from the first iteration.

[0089] If not satisfied The set conditions are then used for a second iteration.

[0090]

[0091] According to Formula 2, also:

[0092]

[0093] Right now:

[0094]

[0095] Substituting Formula 3 into Formula 1 for a second iteration follows the same principle as the first iteration.

[0096]

[0097] This process can be repeated multiple times until the working strain ratio |ε| is reached. y n / γ xy n | The ratio ε that is infinitely close to the design allowable valuead / γ ad In this way, the stiffness trimming of the composite material perfectly matches the structural load-bearing characteristics, resulting in relatively high material design efficiency. Therefore, this is the process of matching the mechanical properties of the material with the load-bearing characteristics of the structure. The final E... x n G xy n The optimal ratio is given by the superscripts, which represent the number of iterations. The calculation principle is to assume that a composite laminate with a certain layup ratio is regarded as a specific single-layer composite material, and to design a composite material with matching performance based on the structural load-bearing characteristics, that is, to design the E and G values ​​of the composite laminate.

[0098] In some alternative implementations, the characteristic values ​​T of different composite laminates in a preset optional layup library are used. m Calculate the applicable boundary characteristic value T for each composite laminate. b The material characteristic value T of composite laminate m The characteristic value T of the composite laminate boundary represents the optimal ratio of axial and shear bearing capacity for this specific proportion of the laminate. b This characterizes the applicable range of axial and shear bearing capacity of this specific ratio of laminate. When the structural bearing characteristic value T is within a certain range of the laminate boundary characteristic value T... b Within the corresponding range, select the boundary feature value T. b The corresponding material characteristic value T m and the corresponding composite laminate materials;

[0099] T m The calculation formula is: Where E 1材 G represents the axial modulus of the composite laminate material. 1材 The shear modulus of the composite laminate material; material characteristic value T b =kT m Where k is a correction factor. The principle is that in structural design, due to various reasons (including process constraints, design experience, and experimental accumulation), designers often choose commonly used ply ratios from the ply library, but they are unsure how to select a ply ratio from the commonly used ply library that matches the load-bearing characteristics. To solve this problem, laminates with specific ply ratios are considered as anisotropic materials. Material characteristic values ​​for the laminate are defined, and boundary critical values, i.e., boundary characteristic values ​​T, are established. b A correction factor is introduced. The zoning critical value divides the laminate of each ply ratio into an interval in which the load-bearing characteristic value |T| applies.

[0100] In some alternative implementations, the initial axial strain ε x, with shear strain γ xy It was obtained through finite element analysis.

[0101] In some alternative embodiments, the load-bearing characteristic value T and the material characteristic value T m With boundary eigenvalue T b Range matching, specifically, involves selecting the material feature value T that is closest to |T|. m And less than the material's characteristic value T m The corresponding boundary eigenvalue T b .

[0102] The advantages of this application include: establishing the relationship between the mechanical requirements and materials of the aircraft ply structure by proposing load-bearing characteristic values ​​and material characteristic values; and making the plying of the aircraft ply structure more targeted through digital quantitative characterization. When plying the aircraft ply structure, the corresponding plying method can be selected according to the different loads of different parts of the aircraft ply structure; and realizing the correlation optimization between load-bearing characteristic values ​​and the ply ratio of laminates to match the optimal ply.

[0103] Example: Optimization of the skin ply ratio design for a certain type of aircraft wing panel

[0104] 1. Example of defining a layered library

[0105] The selected composite materials are shown in Table 1:

[0106] Table 1 Mechanical properties of composite prepregs

[0107]

[0108] Commonly used ply ratios are shown in Table 2 (other ply ratios and angles can also be added from the ply library). The axial modulus E of each ply can be calculated based on the ratio. X E y Poisson's ratio u yx Shear modulus G XY According to the formula: material characteristic value Characteristic critical value T b =T m ×K calculates T for each ply. m T b The results are shown in Table 2.

[0109] Since the E and G values ​​are different for each layup ratio, there is no common benchmark for comparing their performance. Therefore, a correction factor K is introduced into the calculation of the critical value. Using material A1 as the benchmark, the initial critical value of material A2 (which is also the cutoff critical value of material A1) is calculated, with K = G. xy (A2) / G xy(A1); Using material A2 as a benchmark, calculate the initial critical value of material A3 (which is also the cutoff critical value of material A2), and take K = G. xy (A3) / G xy (A2), and so on. Since material A4 has the largest Tm value in the selected layup library, its partitioning threshold is set to ∞, which means that for load-bearing characteristic values ​​exceeding 5.54, the optimal match can only be made using material A4.

[0110] Table 2 Composite Laminate Layup Library

[0111]

[0112] (1) Calculate the load-bearing characteristic value

[0113] The axial stress ε of the element is extracted from the finite element calculation results. x Shear stress γ xy ,calculate Selecting a subset of units, the bearing characteristic value of each unit was obtained, and a matching ply ratio was selected according to the characteristic partition. The results are shown in Table 3. Table 3 shows that T is negative in all cases. According to the definition, the skin bearing shear load in this region is determined by partitioning according to the absolute value of the characteristic value. Regions marked with $ are preferably divided into 4 / 5 / 1 ratios, regions marked with * are preferably divided into 5 / 4 / 1 ratios, and regions marked with @ are preferably divided into 6 / 3 / 1 ratios.

[0114] Table 3 Skin Feature Zoning

[0115]

[0116]

[0117] 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 optimizing the layup ratio of composite laminates, characterized in that, include: Step S1: In the aircraft design process, the design load and aircraft structure layout are determined, and the design allowable value of the composite material used in the aircraft is determined, including the axial allowable strain ε ad , shear allowable strain γ ad ; Step S2: Based on the external loads and specific layout of the aircraft structure, calculate the load-bearing characteristic value T of the aircraft structure to be laid, wherein the load-bearing characteristic value T characterizes the trend characteristics of the axial load and shear load of the aircraft structure to be laid; Step S3: Obtain a plurality of composite material laminates with preset proportions in the ply library, and calculate material characteristic values T of the plurality of composite material laminates with preset proportions m ; Step S4: selecting a material characteristic value T matching the load bearing characteristic value T based on the structure bearing characteristic value T m of the composite laminate; Step S21: presetting an initial elastic modulus E x , an initial shear modulus G xy ; Step S22: based on the initial elastic modulus E x , the initial shear modulus G xy , the initial axial strain ε x , and the shear strain γ xy ; Step S23: through iteration of the elastic modulus E x , the shear modulus G xy , the ratio |ε x / γ xy of the axial strain ε x and the shear strain γ xy is made close to the ratio ε ad / γ ad of the allowable strain ε ad and the allowable stress γ ad , and the elastic modulus E x and the shear modulus G n after iteration are obtained xy n ; Step S24: Based on the iterated elastic modulus E x n With the shear modulus G xy n , the bearing eigenvalue T is calculated.

2. The method for optimizing the layup ratio of composite laminates as described in claim 1, characterized in that, The axial strain ε x The shear strain γ xy The method comprises the following steps: dividing the aircraft layer structure to be laid into multiple units, calculating the bearing characteristic value T of each unit respectively, and selecting a matching composite laminate based on the bearing characteristic value T of each unit respectively.

3. The method for optimizing the layup ratio of composite laminates as described in claim 1, characterized in that, Based on the different composite laminate material characteristic value T in the preset optional layer library m , calculate the boundary characteristic value T of each composite laminate b , the composite laminate material characteristic value T m Characterizes the optimal ratio of axial and shear bearing capacity of this specific ratio laminate, the boundary characteristic value T of the composite laminate b Characterizes the critical value of the axial and shear bearing capacity of this specific ratio laminate, when the structure bearing characteristic value T is in the range corresponding to the boundary characteristic value T b , select the material characteristic value T b Corresponding to the boundary characteristic value T m And the corresponding composite laminate material.

4. The method for optimizing the layup ratio of composite laminates as described in claim 1, characterized in that, The specific formula for calculating the load-bearing characteristic value T is as follows: When ε x n If the strain is tensile, then T = |T| is the tensile-shear characteristic value; if ε x n If the strain is compressive, then T = -|T| is the characteristic value of compressive shear. wherein the iterated axial strain ε x n the iterated shear strain γ xy n the iterated axial stress σ x n the iterated Poisson's ratio μ yx n the iterated normal stress σ y n the iterated shear stress τ xy n .

5. The method for optimizing the layup ratio of composite laminates as described in claim 3, characterized in that, Boundary characteristic value T b = kT m ; where k is a correction factor.

6. The method for optimizing the layup ratio of composite laminates as described in claim 1, characterized in that, T m The calculation formula is: Where E 1材 G represents the axial modulus of the composite laminate material. 1材 This is the shear modulus of the composite laminate material.

7. The method for optimizing the layup ratio of composite laminates as described in claim 1, characterized in that, Initial axial strain ε x With shear strain γ xy Calculated by finite element analysis.

8. The method for optimizing the layup ratio of composite laminates as described in claim 3, characterized in that, The bearing characteristic value T and the material characteristic value T m The boundary characteristic value T b Range matching, and the specific matching method is: selecting a material characteristic value T m Closest to |T|, and less than the material characteristic value T m The corresponding boundary characteristic value T b .

Images