Aircraft typical structure multi-axial fatigue life prediction method based on equivalent stress

Abstract

The invention belongs to the field of airplane strength fatigue life prediction, and particularly relates to an equivalent stress-based airplane typical structure multi-axial fatigue life prediction method. The method comprises the steps of determining a typical aircraft high-lock connection structure test piece; determining a fatigue load spectrum; constructing a finite element model of the typical high-lock connection structure of the aircraft; determining a dangerous part of the airplane typical high-lock connection structure test piece in finite element analysis; determining the maximum principal stress at the dangerous part of the typical airplane high-lock connection structure test piece in finite element analysis; determining the fatigue life of the test piece; fitting S-N curve data under different stress ratios according to the fatigue life of the test piece and the load spectrum stress ratio, and determining a corrected S-N curve; determining an equivalent stress-life equation of the typical high-lock connection structure of the aircraft, and substituting the maximum principal stress and fitting parameters of the corrected S-N curve into the equivalent stress-life equation to calculate and obtain a fatigue prediction life; and correcting the fatigue prediction life according to a test result.

Description

Technical Field

[0001] This application belongs to the field of aircraft strength fatigue life prediction, and specifically relates to a method for predicting the multi-axis fatigue life of typical aircraft structures based on equivalent stress. Background Technology

[0002] The modified DFR method is mainly used for durability assessment of civil aircraft both domestically and internationally. However, the modified DFR method has limitations in calculating fatigue life under multi-axis loading conditions. Analysis of test specimen life using the modified DFR method reveals that the calculation results are overly conservative and unsuitable for predicting the life of fatigue test specimens under multi-axis loading. Therefore, proposing a new life model to more accurately predict fatigue life is of great significance.

[0003] 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

[0004] The purpose of this application is to provide a method for predicting the multiaxial fatigue life of typical aircraft structures based on equivalent stress, so as to solve at least one problem existing in the prior art.

[0005] The technical solution of this application is:

[0006] A method for predicting the multiaxial fatigue life of typical aircraft structures based on equivalent stress includes:

[0007] S1. Determine the test specimen for a typical high-lock connection structure of an aircraft;

[0008] S2. Determine the fatigue load spectrum;

[0009] S3. Construct a finite element model of a typical high-lock connection structure for aircraft;

[0010] S4. Identify the critical locations of typical high-lock connection structure test specimens for aircraft in finite element analysis;

[0011] S5. Determine the maximum principal stress at the critical location of a typical high-lock connection structure test piece for aircraft in finite element analysis.

[0012] S6. Based on the test load spectrum, load the typical high-lock connection structure test piece of the aircraft until the test piece fails to determine the fatigue life of the test piece;

[0013] S7. Fit the SN curve data under different stress ratios based on the fatigue life and load spectrum stress ratio of the test piece to determine the corrected SN curve;

[0014] S8. Determine the equivalent stress-life equation for a typical high-strength connection structure of an aircraft, and substitute the maximum principal stress and the fitting parameters of the modified SN curve into the equivalent stress-life equation to calculate the fatigue prediction life.

[0015] S9. Correct the fatigue prediction life based on the test results.

[0016] In at least one embodiment of this application, in S1, the typical high-lock connection structure test specimen for aircraft includes a tensile-shear test specimen and a bidirectional tensile test specimen.

[0017] In at least one embodiment of this application, in S2, the fatigue load spectrum is a sinusoidal constant amplitude spectrum.

[0018] In at least one embodiment of this application, in S3, a full-size finite element model of the typical high-lock connection structure of the aircraft is constructed using ABAQUS software.

[0019] In at least one embodiment of this application, in S3, the modeling principles of the finite element model of the typical high-lock connection structure of the aircraft include:

[0020] Consider the interference fit and preload at the high-lock bolt;

[0021] The mesh was refined, and C3D8I elements were used for calculation;

[0022] The material property is set to elastic-plastic.

[0023] Load simulation testing machine chuck loading method;

[0024] The boundary conditions are set to allow displacement only in the direction of the load.

[0025] In at least one embodiment of this application, in S4, the dangerous part is determined based on the stress distribution in the finite element model of the typical high-lock connection structure of the aircraft, combined with the loading method and force transmission path.

[0026] In at least one embodiment of this application, in S4, the dangerous part is the edge of the hole in the test area of ​​a typical high-lock connection structure test piece for aircraft.

[0027] In at least one embodiment of this application, in S5, the maximum principal stress is used to evaluate structural safety in finite element analysis, applicable to brittle materials or failure modes dominated by tension and compression.

[0028] In at least one embodiment of this application, S7 includes:

[0029] Obtain the SN curve of the material standard part. The SN curve is plotted with the alternating stress of the material standard part on the vertical axis and the number of cycles on the horizontal axis.

[0030] The SN curve data under different stress ratios were fitted based on the fatigue life and load spectrum stress ratio of the test specimens to determine the corrected SN curve.

[0031] In at least one embodiment of this application, in S8, the equivalent stress-life equation for a typical high-lock connection structure of an aircraft is:

[0032] ;

[0033] ;

[0034] Where, N f For fatigue life prediction, A1, A2, and A3 are the fitting parameters for the modified SN curve, and S eq For equivalent stress, S max R is the maximum principal stress, and R is the stress ratio.

[0035] In at least one embodiment of this application, in S9, the fatigue prediction life is corrected based on test results:

[0036] N=λN f ;

[0037] Where N is the fatigue prediction life correction value, and λ is the correction coefficient.

[0038] The invention has at least the following beneficial technical effects:

[0039] The method for predicting the multi-axis fatigue life of typical aircraft structures based on equivalent stress in this application, which is designed for multi-axis loading of aircraft high-lock connection structures, can predict the fatigue life of the connection structure under multi-axis loading by combining the finite element method with fatigue equivalent stress, thus facilitating the prediction of fatigue life of multi-axis loaded structures. Attached Figure Description

[0040] Figure 1 This is an axial view of a tensile-shear test specimen according to one embodiment of this application;

[0041] Figure 2 This is a front view of a tensile-shear test specimen according to one embodiment of this application;

[0042] Figure 3 This is a top view of a tensile-shear test specimen according to one embodiment of this application;

[0043] Figure 4 This is an axial view of a biaxially stressed test specimen according to one embodiment of this application;

[0044] Figure 5 This is a front view of a biaxial tensile test specimen according to one embodiment of this application;

[0045] Figure 6 This is a side view of a biaxial tensile test specimen according to one embodiment of this application;

[0046] Figure 7 This is a spectrum of the critical location of a tension-shear model according to one embodiment of this application;

[0047] Figure 8 This is a spectrum of the critical location cloud map of the tensile-shear model according to one embodiment of this application;

[0048] Figure 9 This is a spectrum of the critical location cloud map of the tensile-shear model according to one embodiment of this application;

[0049] Figure 10 This is a cloud map of the critical part of the bidirectional tensile model spectral load 1 according to one embodiment of this application;

[0050] Figure 11 This is a cloud map of the critical part of the bidirectional tensile model under spectral load 2 according to one embodiment of this application;

[0051] Figure 12 This is a cloud map of the critical location of the bidirectional tensile model under spectral load in one embodiment of this application;

[0052] Figure 13 This is the SN curve of a 1-inch thick 7050-T7451 aluminum alloy flat plate under different stress ratios according to one embodiment of this application;

[0053] Figure 14 This is a modified SN curve of one embodiment 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 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.

[0055] 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.

[0056] The following is in conjunction with the appendix Figures 1 to 14 This application will be described in further detail.

[0057] This application provides a method for predicting the multi-axis fatigue life of typical aircraft structures based on equivalent stress, including the following steps:

[0058] S1. Determine the test specimen for a typical high-lock connection structure of an aircraft;

[0059] S2. Determine the fatigue load spectrum;

[0060] S3. Construct a finite element model of a typical high-lock connection structure for aircraft;

[0061] S4. Identify the critical locations of typical high-lock connection structure test specimens for aircraft in finite element analysis;

[0062] S5. Determine the maximum principal stress at the critical location of a typical high-lock connection structure test piece for aircraft in finite element analysis.

[0063] S6. Based on the test load spectrum, load the typical high-lock connection structure test piece of the aircraft until the test piece fails to determine the fatigue life of the test piece.

[0064] S7. Fit the SN curve data under different stress ratios based on the fatigue life and load spectrum stress ratio of the test piece to determine the corrected SN curve;

[0065] S8. Determine the equivalent stress-life equation for a typical high-strength connection structure of an aircraft, and substitute the maximum principal stress and the fitting parameters of the modified SN curve into the equivalent stress-life equation to calculate the fatigue prediction life.

[0066] S9. Correct the fatigue prediction life based on the test results.

[0067] This application presents a method for predicting the multiaxial fatigue life of typical aircraft structures based on equivalent stress. In S1, the typical high-lock connection structure test specimens for aircraft include two types: tensile-shear test specimens and biaxially tensile test specimens. In S2, the fatigue load spectrum of the multiaxially loaded structure is a sinusoidal waveform with constant amplitude. In S3, full-size finite element models of the two types of typical high-lock connection structure test specimens for aircraft are constructed using ABAQUS software. The specific modeling principles are as follows:

[0068] To more realistically represent the stress conditions at the edge of the hole, interference fit and preload at the high-locking bolt are considered;

[0069] The mesh was refined, and C3D8I elements were used for calculation;

[0070] The material property is set to elastic-plastic.

[0071] Load simulation testing machine chuck loading method;

[0072] The boundary conditions are set to allow displacement only in the direction of the load.

[0073] In S4, the critical location is determined based on the stress distribution in the finite element model of the typical high-strength interlocking connection structure of the aircraft, combined with the loading method and force transmission path. The critical location is usually the edge of the hole in the test area of ​​the typical high-strength interlocking connection structure test piece of the aircraft. In S5, the maximum principal stress is used to evaluate the structural safety in finite element analysis, and is applicable to brittle materials or failure modes dominated by tension and compression. In S7, the SN curve of the material standard part is obtained. The SN curve is plotted with the alternating stress of the material standard part as the ordinate and the number of cycles as the abscissa. The SN curve data under different stress ratios are fitted according to the fatigue life of the test piece and the stress ratio of the load spectrum to determine the corrected SN curve. The stress ratio R is the ratio of the minimum stress to the maximum stress under alternating stress. In S8, the equivalent stress-life equation of the typical high-strength interlocking connection structure of the aircraft is:

[0074] ;

[0075] ;

[0076] Where, N f For fatigue life prediction, A1, A2, and A3 are the fitting parameters for the modified SN curve, and S eq For equivalent stress, S max R is the maximum principal stress, and R is the stress ratio.

[0077] In S9, considering the effects of factors such as the assembly stress of the test piece and the error in the fitting results, the fatigue prediction life is corrected based on the test results:

[0078] N=λN f ;

[0079] Where N is the fatigue prediction life correction value, and λ is the correction coefficient.

[0080] In one embodiment of this application, multiaxial fatigue life prediction based on equivalent stress is proposed for two typical aircraft high-lock connection structures, namely, tensile-shear test specimens and biaxial tensile test specimens. The specific implementation steps of the multiaxial fatigue life prediction method for typical aircraft structures based on equivalent stress according to this application are as follows:

[0081] S1. Determine the test specimen for a typical high-lock connection structure of an aircraft;

[0082] Test specimen I is a tensile-shear biaxial fatigue test specimen, such as Figure 1-3 As shown, the test specimen consists of one flat plate test specimen and four T-shaped test specimens, and is subjected to bidirectional tensile and shear loads.

[0083] Test specimen ⅠⅠ is a biaxial tensile fatigue test specimen, such as Figure 4-6As shown, the test specimen consists of one flat plate specimen and four strip plate specimens, and is subjected to biaxial tensile load.

[0084] The main parameters of the test specimens are shown in Table 1.

[0085] Table 1

[0086]

[0087] S2. Determine the fatigue load spectrum;

[0088] The fatigue load conditions are shown in Table 2.

[0089] Table 2

[0090]

[0091] S3. Construct a finite element model of a typical high-lock connection structure for aircraft;

[0092] To analyze the stress distribution of the tensile-shear test specimen, a full-size finite element model of the specimen was constructed using ABAQUS software. The specific settings are as follows:

[0093] a) To more realistically represent the stress conditions at the hole edge, interference fit and preload at the high-lock bolt are considered. The interference is taken as 1%, and the preload is taken as 7KN. To accurately obtain the stress value at the critical location (stress concentration point), the mesh is refined, and C3D8I elements are used for calculation.

[0094] b) The material properties in the model are set to elastoplastic. The base plate (flat structure test piece) and strip plate (T-shaped structure test piece) are made of 7050-T7451 aluminum alloy with an elastic modulus of 71.7 GPa and a Poisson's ratio of 0.3. The high-locking bolts in the test area are made of TC4 titanium alloy with an elastic modulus of 110 GPa and a Poisson's ratio of 0.3. The loading end pins are made of 42CrMo steel with an elastic modulus of 212 GPa and a Poisson's ratio of 0.28.

[0095] c) X-direction load simulation test machine clamp loading method; Y-direction load is applied to the end face of the clamp.

[0096] d) The boundary conditions are set to allow displacement only at the two end faces of the substrate along the X direction and displacement only at the end faces of the two clamps along the Y direction.

[0097] Finite element simulation of the biaxial tension test specimen was performed using ABAQUS software. The preprocessing process was the same as that of the tension-shear test specimen. The X and Y loads were simulated by the loading method of the testing machine clamp. The boundary conditions were set to allow displacement of the two end faces of the substrate along the X direction and displacement of the end face of the strip along the Y direction.

[0098] S4. Identify the critical locations of typical high-lock connection structure test specimens for aircraft in finite element analysis;

[0099] Analysis of the stress distribution in the finite element model, combined with the loading method and force transmission path, shows that the hole edges in the test area of ​​the substrate of the tensile-shear test specimen and the biaxial tensile test specimen are the areas with the most severe stress concentration. Figure 7-12 As shown.

[0100] S5. Determine the maximum principal stress at the critical location of a typical high-lock connection structure test piece for aircraft in finite element analysis.

[0101] The maximum principal stresses at critical locations for tensile-shear and biaxial tensile test specimens under the three spectral loads are shown in Table 3.

[0102] Table 3

[0103]

[0104] S6. Determine the fatigue life of the test specimen;

[0105] Table 4 shows the fatigue failure life of the tensile-shear test specimens and the biaxial tensile test specimens under three different load ranges.

[0106] Table 4

[0107]

[0108] S7. Fit the SN curve data under different stress ratios based on the fatigue life and load spectrum stress ratio of the test piece to determine the corrected SN curve;

[0109] For 7050-T7451 aluminum alloy, its SN curves under different stress ratios are as follows: Figure 13 As shown. The fatigue life of the test piece and... Figure 13 A corrected SN curve was obtained by fitting the SN curve data with a stress ratio of 0.06. The fitting results are as follows: Figure 14 As shown:

[0110] S8. Calculate the predicted fatigue life;

[0111] For different stress ratios R, the stress-controlled fatigue life can be fitted by the equivalent stress-life equation. Based on the fitting results, the equivalent stress-life equation is obtained as follows:

[0112] ;

[0113] ;

[0114] The fatigue prediction life is calculated based on the equivalent stress-life equation.

[0115] Finally, S9, adjust the fatigue prediction life based on the test results;

[0116] Based on the test life results, the correction factor was confirmed to be 0.80.

[0117] By combining the equivalent stress-life equation with the maximum principal stress at the critical location, the fatigue life of the modified tensile-shear test specimen under three spectral loads can be obtained, as shown in Table 5.

[0118] Table 5

[0119]

[0120] By combining the equivalent stress-life equation with the maximum principal stress at the critical location, the fatigue life of the modified biaxial tensile specimen under three spectral loads can be obtained, as shown in Table 6.

[0121] Table 6

[0122]

[0123] 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 predicting the multiaxial fatigue life of typical aircraft structures based on equivalent stress, characterized in that, include: S1. Determine the test specimen for a typical high-lock connection structure of an aircraft; S2. Determine the fatigue load spectrum; S3. Construct a finite element model of a typical high-lock connection structure for aircraft; S4. Identify the critical locations of typical high-lock connection structure test specimens for aircraft in finite element analysis; S5. Determine the maximum principal stress at the critical location of a typical high-lock connection structure test piece for aircraft in finite element analysis. S6. Based on the test load spectrum, load the typical high-lock connection structure test piece of the aircraft until the test piece fails to determine the fatigue life of the test piece; S7. Fit the SN curve data under different stress ratios based on the fatigue life and load spectrum stress ratio of the test piece to determine the corrected SN curve; S8. Determine the equivalent stress-life equation for a typical high-strength connection structure of an aircraft, and substitute the maximum principal stress and the fitting parameters of the modified SN curve into the equivalent stress-life equation to calculate the fatigue prediction life. S9. Correct the fatigue prediction life based on the test results.

2. The method for predicting the multi-axis fatigue life of typical aircraft structures based on equivalent stress according to claim 1, characterized in that, In S1, the typical high-lock connection structure test specimen for aircraft includes a tensile-shear test specimen and a bidirectional tensile test specimen.

3. The method for predicting the multi-axis fatigue life of typical aircraft structures based on equivalent stress according to claim 2, characterized in that, In S2, the fatigue load spectrum is a sinusoidal waveform with constant amplitude.

4. The method for predicting the multiaxial fatigue life of typical aircraft structures based on equivalent stress according to claim 3, characterized in that, In S3, a full-size finite element model of the typical high-lock connection structure of the aircraft is constructed using ABAQUS software.

5. The method for predicting the multi-axis fatigue life of typical aircraft structures based on equivalent stress according to claim 4, characterized in that, In S3, the modeling principles of the finite element model of the typical high-lock connection structure of the aircraft include: Consider the interference fit and preload at the high-lock bolt; The mesh was refined, and C3D8I elements were used for calculation; The material property is set to elastic-plastic. Load simulation testing machine chuck loading method; The boundary conditions are set to allow displacement only in the direction of the load.

6. The method for predicting the multiaxial fatigue life of typical aircraft structures based on equivalent stress according to claim 5, characterized in that, In S4, the dangerous part is determined based on the stress distribution in the finite element model of the typical high-lock connection structure of the aircraft, combined with the loading method and force transmission path.

7. The method for predicting the multiaxial fatigue life of typical aircraft structures based on equivalent stress according to claim 6, characterized in that, In S4, the dangerous part is the edge of the hole in the test area of ​​a typical high-lock connection structure test piece for aircraft.

8. The method for predicting the multi-axis fatigue life of typical aircraft structures based on equivalent stress according to claim 7, characterized in that, In S5, the maximum principal stress is used to evaluate structural safety in finite element analysis and is applicable to brittle materials or failure modes dominated by tension and compression.

9. The method for predicting the multi-axis fatigue life of typical aircraft structures based on equivalent stress according to claim 8, characterized in that, S7 includes: Obtain the SN curve of the material standard part. The SN curve is plotted with the alternating stress of the material standard part on the vertical axis and the number of cycles on the horizontal axis. The SN curve data under different stress ratios were fitted based on the fatigue life and load spectrum stress ratio of the test specimens to determine the corrected SN curve.

10. The method for predicting the multiaxial fatigue life of typical aircraft structures based on equivalent stress according to claim 9, characterized in that, In S8, the equivalent stress-life equation for a typical high-lock connection structure in an aircraft is: ； ； Where, N f For fatigue life prediction, A1, A2, and A3 are fitting parameters for the modified SN curve, and S... eq For equivalent stress, S max R is the maximum principal stress, and R is the stress ratio.

11. The method for predicting the multiaxial fatigue life of typical aircraft structures based on equivalent stress according to claim 10, characterized in that, In S9, the fatigue prediction life is corrected based on the test results: N=λN f ; Where N is the fatigue prediction life correction value, and λ is the correction coefficient.

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