Fatigue damage tolerance evaluation method for aerospace power component in vibration environment

Abstract

The invention relates to a fatigue damage tolerance evaluation method for a spaceflight power component in a vibration environment. The method is mainly used for solving the technical problem that the residual service life of a typical component of a spaceflight power system in a complex vibration load environment is difficult to accurately predict by an existing life prediction method. The method comprises the following steps: modeling a to-be-evaluated component and an initial crack on the to-be-evaluated component, performing transient dynamic analysis on a finite element model to obtain dynamic stress-time history data of the crack, and calculating the dynamic stress-time history data by adopting a stress extrapolation method to obtain stress intensity factor K-time history data of the crack; the time-domain variable-amplitude load is introduced into fracture mechanical modeling. And performing cyclic statistical processing on the stress intensity factor K-time history data by adopting a rain flow counting method to obtain a three-variable rain flow matrix of a stress intensity factor range delta K-cyclic ratio R-cyclic index n, and performing high-precision prediction on the residual life of the to-be-evaluated component.

Description

TECHNICAL FIELD

[0001] The present application relates to a fatigue damage tolerance evaluation method of a spaceflight power component, in particular to a fatigue damage tolerance evaluation method of a spaceflight power component under a vibration environment. BACKGROUND

[0002] In the field of aerospace, spacecraft components are subjected to complex service environments such as high pressure, extreme temperature difference and strong vibration, and parts are prone to fatigue failure, which in turn leads to crack initiation and propagation. Especially during the launch process, on-orbit operation and the process of returning to the atmosphere, the service conditions of spacecraft components are deteriorating, which easily leads to fatigue fracture failure of the components. In the field of aerospace engineering, which is highly complex and has extremely strict safety requirements, when a component initiates a fatigue crack, the common treatment strategy is to replace the part directly. Although this method can quickly restore the integrity and reliability of the system, the overly conservative maintenance strategy leads to high economic costs. Therefore, it is of great theoretical and practical significance to establish a residual life analysis method for typical components of a spaceflight power system with damage under a vibration environment for the realization of an economic and efficient maintenance strategy of a spacecraft structure.

[0003] In the field of aerospace, the requirements for structural integrity of spacecraft under high-frequency vibration environment are far beyond those of general structures. Currently, the structural integrity assessment of typical components of a spaceflight power system mainly adopts traditional infinite life criteria and safe life criteria. However, when assessing the safety of structures with defects and damage, these two criteria cannot accurately predict and evaluate the influence of defects on the performance of the structure, resulting in insufficient accuracy of life prediction. SUMMARY

[0004] The purpose of the present application is to solve the technical problem that the existing life prediction method cannot accurately predict the residual service life of typical components of a spaceflight power system under complex vibration load environment, and to propose a fatigue damage tolerance evaluation method of a spaceflight power component under a vibration environment.

[0005] To solve the above technical problems, the technical solution provided by the present application is as follows: A fatigue damage tolerance evaluation method of a spaceflight power component under a vibration environment, characterized in that it comprises the following steps: S1, obtaining necessary parameters and a load spectrum of the component to be evaluated, cleaning and verifying the necessary parameters and the load spectrum, and removing outliers; the load spectrum is a time-domain variable-amplitude vibration load; S2, establishing a finite element model of the component to be evaluated according to the necessary parameters, and modeling a crack in the finite element model to obtain an initial crack length a0; S3, meshing the finite element model so that the mesh size and density at the crack and the stress extrapolation zone are greater than those in other regions; S4, import the finite element model into the CAE software to give the material properties of the finite element model, so that each mesh has corresponding material parameters; according to the actual working condition of the component to be evaluated, the boundary conditions are applied; S5, based on the load spectrum, material parameters and boundary conditions, the transient dynamics analysis is carried out on the finite element model, and the dynamic stress-time history data of the crack are obtained; S6, the stress distribution of the nodes near the crack tip is extrapolated by the stress extrapolation method, the dynamic stress-time history data of the crack are calculated, and the stress intensity factor K-time history data of the crack are obtained; S7, the stress intensity factor K-time history data are cyclically statistically processed by the rainflow counting method, and a three-variable rainflow matrix of the stress intensity factor range ΔK-cycle ratio R-cycle number n is obtained; one row of parameters in the three-variable rainflow matrix is one type of cycle, and there are I rows in total; S8, the parameters in the three-variable rainflow matrix are read row by row, and the cycle step-by-step integration method is used to simulate the crack length after each type of cycle, and when the crack length reaches the predetermined failure criterion, the total cycle number is calculated as the predicted remaining life, and the fatigue damage tolerance evaluation is completed.

[0006] Further, step S8 specifically comprises: S81, the crack increment (da / dN)_i in each type of cycle is calculated by the Paris equation, wherein a represents the crack length, and N represents the cycle number; the average fatigue crack propagation rate (da / dN) is obtained by averaging the I crack increments (da / dN)_i. mean S82, the stress intensity factor range ΔK1, cycle ratio R1 and cycle number n1 in the first row of parameters in the three-variable rainflow matrix are read, and the first type of cycle is carried out; the crack length after the first type of cycle is expanded is preset as a1; S83, the crack propagation increment Δa1 under the first type of cycle is calculated, Δa = a1− a0; The crack propagation life ΔN1 under the first type of cycle is calculated by the cycle step-by-step integration method: ; The predicted remaining life is N1=ΔN1; S84, according to the method of step S82, the stress intensity factor range ΔKi, cycle ratio Ri and cycle number ni in the i-th row of parameters in the three-variable rainflow matrix are read row by row, and the i-th type of cycle is carried out; the crack length after the i-th type of cycle is expanded is preset as ai, and the reading is stopped until the crack length ai reaches the critical size; S85, the crack propagation increment Δai under the i-th type of cycle is calculated, Δa = ai− a(i-1); ​The crack propagation life ΔNi under the i-th type of cycle is calculated using the cyclic successive integration method: ; Predict the remaining life Ni = N(i-1) + ΔNi, and complete the fatigue damage tolerance evaluation.

[0007] Furthermore, step S5 specifically includes: S51. Perform modal analysis on the finite element model to obtain the natural frequencies and mode shapes of the finite element model; S52. Using the natural frequency and mode shape as initial conditions, perform transient dynamic analysis on the finite element model and calculate the dynamic response of the finite element model under the load spectrum. S53. Extract dynamic stress-time history data of crack front nodes from dynamic response.

[0008] Furthermore, in step S1, the necessary parameters include geometric dimensions, material properties, and the geometric shape and spatial location of the crack; The load spectrum includes the magnitude, direction, and frequency of the load borne by the component to be evaluated.

[0009] Furthermore, in step S2, the specific steps for modeling the crack in the finite element model are as follows: S21. Predict and calculate the crack propagation direction and possible paths; S22. Model the crack based on the geometric shape, spatial location, propagation direction, and possible paths of the crack in the necessary parameters.

[0010] Furthermore, step S6 specifically includes: S61. Select h nodes at the crack tip and obtain the dynamic stress values ​​σ at the h nodes based on the dynamic stress-time history data. i , 5≤h≤20; S62. Using the stress extrapolation method, based on the dynamic stress value σ at the node... i Calculate the corresponding stress intensity factor K i ; S63. Construct data point pairs (K) for each node. i r i ), where r i This represents the radial distance of the i-th node from the crack tip; S64. Use the least squares method to process the data point pairs (K) of each node. i r i Perform linear regression fitting, and calculate the slope A and intercept B of the fitted line: ; ; S65. Approximate expression for calculating the radial distribution of stress intensity factor. The stress intensity factor K-time history data for cracks: .

[0011] Compared with the prior art, the beneficial effects of the present invention are: 1. This invention discloses a fatigue damage tolerance evaluation method for aerospace propulsion components under vibration conditions. The method involves modeling the component to be evaluated and its initial cracks, performing transient dynamic analysis on the finite element model to obtain dynamic stress-time history data of the crack, and calculating the stress intensity factor K-time history data of the crack using stress extrapolation. This allows for the introduction of time-domain variable amplitude loads into fracture mechanics modeling. Rainflow counting is used to perform cyclic statistical processing on the stress intensity factor K-time history data, obtaining a three-variable rainflow matrix consisting of the stress intensity factor range ΔK, cycle ratio R, and cycle number n, enabling high-precision prediction of the remaining life of the component to be evaluated.

[0012] 2. The present invention provides a fatigue damage tolerance evaluation method for aerospace power components under vibration environment. The vibration load is introduced into fracture mechanics analysis as a high-precision time-domain signal, providing details of the vibration load change over time for the fracture mechanics model, so that the finite element model can more comprehensively and realistically reflect the service environment of the component under complex dynamic load.

[0013] 3. This invention provides a fatigue damage tolerance evaluation method for aerospace propulsion components under vibration conditions. It employs a stress extrapolation method, indirectly calculating the stress intensity factor K by extrapolating the stress distribution near the crack tip, thereby obtaining the stress intensity factor-time history data at the crack tip. This method can accurately capture the crack propagation characteristics under complex time-varying loads.

[0014] 4. The present invention provides a fatigue damage tolerance evaluation method for aerospace propulsion components under vibration environment. It combines the Paris equation and the iterative step-by-step integration method to predict the remaining life of typical aerospace propulsion system components under vibration fatigue load, thus solving the problem of remaining life evaluation for defective components. Attached Figure Description

[0015] Figure 1 This is a flowchart of an embodiment of the fatigue damage tolerance evaluation method for aerospace power components under vibration environment according to the present invention; Figure 2 This is a schematic diagram of the finite element model in an embodiment of the fatigue damage tolerance evaluation method for aerospace power components under vibration environment according to the present invention. In the figure, a is the finite element model, and b is the crack; Figure 3This is a curve of stress intensity factor K-time history data in an embodiment of the fatigue damage tolerance evaluation method for aerospace power components under vibration environment according to the present invention. Figure 4 This is a three-variable rainflow matrix diagram in an embodiment of the fatigue damage tolerance evaluation method for aerospace power components under vibration environment of the present invention; Figure 5 This is a diagram showing the predicted remaining life in an embodiment of the fatigue damage tolerance evaluation method for aerospace power components under vibration environment according to the present invention. Detailed Implementation

[0016] The present invention will be further described below with reference to the accompanying drawings and embodiments.

[0017] This invention provides a method for evaluating the fatigue damage tolerance of aerospace power components under vibration conditions, such as... Figure 1 As shown, it includes the following steps: S1. Obtain the necessary parameters and load spectrum of the component to be evaluated. Clean and verify the necessary parameters and load spectrum to remove outliers. Obtaining the necessary parameters and load spectrum is the basic work for life prediction. Ensure the accuracy and reliability of the data to provide accurate data input for subsequent fatigue analysis. Essential parameters include geometric dimensions, material properties (such as elastic modulus, Poisson's ratio, material density, etc.), and the geometric shape and spatial location of the crack (such as crack length, width, depth, direction, etc.). The load spectrum includes the magnitude, direction, and frequency of the loads (such as static loads, dynamic loads, and impact loads) borne by the component under evaluation. Comprehensive load spectrum information is obtained through methods such as field measurements, historical data analysis, or simulation.

[0018] S2. Based on the necessary parameters and defined geometric dimensions and material properties, establish a finite element model of the component to be evaluated, and model the crack within the finite element model to obtain the initial crack length a0, such as... Figure 2 As shown.

[0019] To ensure accurate crack implantation, it is necessary to consider not only the initial morphology and location of the crack, but also its propagation direction and path. This provides a basis for subsequent crack propagation analysis and ensures that the crack is accurately implanted in the finite element model.

[0020] S21. Predict and calculate the crack propagation direction and possible paths; S22. Model the crack based on the geometric shape, spatial location, propagation direction, and possible paths of the crack in the necessary parameters.

[0021] When modeling, pay attention to the detailed features of the structure, such as welds and bolted connections, as these features have a significant impact on the strength and fatigue performance of the structure. Ensure that the model accurately reflects the actual shape, size, and material properties of the object under study.

[0022] S3. Mesh quality directly affects the accuracy of the analysis results and computational efficiency. Therefore, it is necessary to carefully set the mesh size and shape, and to refine the mesh locally at crack locations and stress extrapolation zones. At the same time, the mesh can be appropriately sparsed away from cracks and stress extrapolation zones to reduce the amount of computation and minimize simulation time while ensuring computational accuracy.

[0023] Use an appropriate mesh type (such as a hexahedral mesh) to mesh the finite element model, and perform local mesh refinement in the stress extrapolation zone at the crack tip to make the mesh size and density at the crack and stress extrapolation zone larger than in other areas, thereby improving the calculation accuracy. S4. Import the finite element model into CAE (Computer-aided Engineering) software, such as Ansys or Abaqus. Assign material properties to the finite element model so that each mesh has corresponding material parameters. Apply boundary conditions according to the actual working conditions of the component to be evaluated, such as fixed constraints, displacement constraints, and loads. For complex working conditions, it is necessary to consider the combination and variation of multiple boundary conditions. S5. Based on the load spectrum, material parameters and boundary conditions, perform transient dynamic analysis on the finite element model to obtain the dynamic stress-time history data of the crack; S51. Perform modal analysis on the finite element model to obtain the natural frequencies and mode shapes of the finite element model; S52. Using the natural frequency and mode shape as initial conditions, perform transient dynamic analysis on the finite element model and calculate the dynamic response of the finite element model under the load spectrum. S53. Extract dynamic stress-time history data of crack front nodes from dynamic response.

[0024] S6. Using the stress extrapolation method, based on the theoretical model of the stress field near the crack tip, the stress distribution at the nodes near the crack tip is extrapolated. The dynamic stress-time history data of the crack is calculated to obtain the stress intensity factor K-time history data of the crack, such as... Figure 3 As shown; The stress intensity factor K is a physical quantity that measures the degree of stress concentration at the crack tip and is used to assess the stress concentration near the crack tip.

[0025] Stress extrapolation is a method used in fracture mechanics to estimate the stress intensity factor K at the crack tip. Based on a theoretical model of the stress field near the crack tip, this method indirectly calculates the stress intensity factor K by extrapolating the stress distribution near the crack tip, thereby obtaining stress intensity factor-time history data at the crack tip. Specifically: S61. Select h nodes at the crack tip and obtain the dynamic stress values ​​σ at the h nodes based on the dynamic stress-time history data. i , 5≤h≤20; S62. Using the stress extrapolation method, based on the dynamic stress value σ at the node... i Calculate the corresponding stress intensity factor K i ; S63. Construct data point pairs (K) for each node. i r i ), where r i This represents the radial distance of the i-th node from the crack tip; S64. Use the least squares method to process the data point pairs (K) of each node. i r i Perform linear regression fitting, and calculate the slope A and intercept B of the fitted line: ; ; S65. Approximate expression for calculating the radial distribution of stress intensity factor. The stress intensity factor K-time history data for cracks: .

[0026] S7. The stress intensity factor K-time history data is statistically processed using the rainflow counting method to obtain a three-variable rainflow matrix consisting of the stress intensity factor range ΔK, the cycle ratio R, and the cycle number n. Figure 4 As shown; one row of parameters in the three-variable rainflow matrix is ​​a type of cycle, with a total of I rows; Rainflow counting is an algorithm used to extract fatigue load cycles from random load histories. By statistically analyzing random sequences using rainflow counting, the range of stress intensity factors (i.e., the difference between the maximum and minimum stress intensity factors) for each cycle can be determined.

[0027] S8. Read the parameters in the three-variable rainflow matrix line by line, and use the iterative step-by-step integration method to simulate the crack length after crack propagation for each type of cycle. When the crack length reaches the predetermined failure criterion, calculate the total number of cycles as the predicted remaining life to complete the fatigue damage tolerance evaluation. Figure 5As shown in the figure, when the crack length increases from 0.1 mm to 3.0 mm in depth, the standard component still has 4.4 × 10⁻⁶ cycles remaining. 8 This value can be interpreted as follows: under a given load spectrum, the component experiences approximately 4.4 × 10⁻⁶ load "block spectrum" cycles. 8 At this point, the crack will propagate from the initial 0.1 mm to 3.0 mm. Specifically: S81. Calculate the crack increment (da / dN)_i in each type of cycle using the Paris equation, where a represents the crack length and N represents the number of cycles; calculate the average fatigue crack propagation rate (da / dN) by averaging the I crack increments (da / dN). mean ; S82. Read the first row of parameters in the three-variable rainflow matrix: stress intensity factor range ΔK1, cycle ratio R1, and cycle number n1, and perform the first type of cycle; preset the crack length after the first type of cycle expansion to a1. S83. Calculate the crack propagation increment Δa1 under the first type of cycle, Δa = a1− a0; The crack propagation life ΔN1 under the first type of cycle was calculated using the cyclic stepwise integration method: ; The predicted remaining lifetime is N1 = ΔN1; S84. Following the method in step S82, read the stress intensity factor range ΔKi, cycle ratio Ri, and cycle number ni of the i-th row parameters in the three-variable rainflow matrix line by line, and perform the i-th type of cycle; preset the crack length after the expansion of the i-th type of cycle to ai, until the crack length ai reaches the critical size, and stop reading; S85. Calculate the crack propagation increment Δai under the i-th type of cycle, Δa = ai− a(i-1); The crack propagation life ΔNi under the i-th type of cycle is calculated using the cyclic successive integration method: ; Predict the remaining life Ni = N(i-1) + ΔNi, and complete the fatigue damage tolerance evaluation.

[0028] The Paris equation is typically expressed as a power function relationship between the crack propagation rate da / dN and the stress intensity factor range ΔK. By fitting experimental data to obtain coefficients and exponents, the crack propagation rate and remaining life under different stress conditions can be predicted.

[0029] By combining the Paris equation and the iterative integration method, the remaining life of a structure under vibration fatigue can be systematically predicted, providing an important reference for engineering practice.

[0030] This invention introduces time-domain loads into damage tolerance assessment based on fracture mechanics and proposes a method for predicting the remaining life of typical aerospace propulsion structures under complex vibration loads. It achieves accurate prediction of the service life of damaged structural components, and has practical application significance and important guiding role for the safe service of typical components in the aerospace field.

Claims

1. A method for evaluating the fatigue damage tolerance of aerospace power components under vibration environment, characterized in that, Includes the following steps: S1. Obtain the necessary parameters and load spectrum of the component to be evaluated, clean and verify the necessary parameters and load spectrum, and remove outliers; The load spectrum is a time-domain amplitude-variable vibration load. S2. Based on the necessary parameters, establish a finite element model of the component to be evaluated, and model the crack in the finite element model to obtain the initial crack length a0. S3. Mesh the finite element model so that the mesh size and density at the crack and stress extrapolation zones are larger than those in other areas; S4. Import the finite element model into the CAE software and assign material properties to the finite element model so that each mesh has the corresponding material parameters. Apply boundary conditions based on the actual working conditions of the component to be evaluated; S5. Based on the load spectrum, material parameters and boundary conditions, perform transient dynamic analysis on the finite element model to obtain the dynamic stress-time history data of the crack; S6. The stress distribution of nodes near the crack tip is extrapolated using the stress extrapolation method. The dynamic stress-time history data of the crack is calculated to obtain the stress intensity factor K-time history data of the crack. S7. The stress intensity factor K-time history data is subjected to cyclic statistical processing using the rainflow counting method to obtain a three-variable rainflow matrix with stress intensity factor range ΔK-cycle ratio R-cycle number n. One row of parameters in the three-variable rainflow matrix is ​​a type of cycle, with a total of I rows; S8. Read the parameters in the three-variable rainflow matrix line by line, and use the iterative step-by-step integration method to simulate the crack length after crack propagation for each type of cycle. When the crack length reaches the predetermined failure criterion, calculate the total number of cycles as the predicted remaining life and complete the fatigue damage tolerance evaluation.

2. The fatigue damage tolerance evaluation method for aerospace power components under vibration environment according to claim 1, characterized in that, Step S8 is as follows: S81. Calculate the crack increment (da / dN)_i in each type of cycle using the Paris equation, where a represents the crack length and N represents the number of cycles; calculate the average fatigue crack propagation rate (da / dN) by averaging the I crack increments (da / dN). mean ; S82. Read the first row of parameters in the three-variable rainflow matrix: stress intensity factor range ΔK1, cycle ratio R1, and cycle number n1, and perform the first type of cycle; preset the crack length after the first type of cycle expansion to a1. S83. Calculate the crack propagation increment Δa1 under the first type of cycle, Δa = a1− a0; The crack propagation life ΔN1 under the first type of cycle was calculated using the cyclic successive integration method: ； The predicted remaining lifetime is N1 = ΔN1; S84. Following the method in step S82, read the stress intensity factor range ΔKi, cycle ratio Ri, and cycle number ni of the i-th row parameters in the three-variable rainflow matrix line by line, and perform the i-th type of cycle. The crack length after the i-th type of cyclic expansion is preset to ai, and reading stops when the crack length ai reaches the critical size; S85. Calculate the crack propagation increment Δai under the i-th type of cycle, Δa = ai− a(i-1); The crack propagation life ΔNi under the i-th type of cycle is calculated using the cyclic successive integration method: ； Predict the remaining life Ni = N(i-1) + ΔNi, and complete the fatigue damage tolerance evaluation.

3. The fatigue damage tolerance evaluation method for aerospace power components under vibration environment according to claim 2, characterized in that, Step S5 is as follows: S51. Perform modal analysis on the finite element model to obtain the natural frequencies and mode shapes of the finite element model; S52. Using the natural frequency and mode shape as initial conditions, perform transient dynamic analysis on the finite element model and calculate the dynamic response of the finite element model under the load spectrum. S53. Extract dynamic stress-time history data of crack front nodes from dynamic response.

4. The fatigue damage tolerance evaluation method for aerospace power components under vibration environment according to claim 3, characterized in that: In step S1, the necessary parameters include geometric dimensions, material properties, and the geometric shape and spatial location of the crack; The load spectrum includes the magnitude, direction, and frequency of the load borne by the component to be evaluated.

5. The fatigue damage tolerance evaluation method for aerospace power components under vibration environment according to claim 4, characterized in that, In step S2, the specific steps for modeling the crack in the finite element model are as follows: S21. Predict and calculate the crack propagation direction and possible paths; S22. Model the crack based on the geometric shape, spatial location, propagation direction, and possible paths of the crack in the necessary parameters.

6. The fatigue damage tolerance evaluation method for aerospace power components under vibration environment according to claim 5, characterized in that, Step S6 is as follows: S61. Select h nodes at the crack tip and obtain the dynamic stress values ​​σ at the h nodes based on the dynamic stress-time history data. i , 5≤h≤20; S62. Using the stress extrapolation method, based on the dynamic stress value σ at the node... i Calculate the corresponding stress intensity factor K i ; S63. Construct data point pairs (K) for each node. i r i ), where r i This represents the radial distance of the i-th node from the crack tip; S64. Use the least squares method to process the data point pairs (K) of each node. i r i Perform linear regression fitting, and calculate the slope A and intercept B of the fitted line: ； ； S65. Approximate expression for calculating the radial distribution of stress intensity factor. The stress intensity factor K-time history data for cracks: 。

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