A low-load truncation method and system for hybrid spectrum fatigue test load spectrum

Abstract

The application discloses a low-load cutting method and system for a mixed structure fatigue test load spectrum, and relates to the technical field of aircraft structure fatigue test, and comprises the following steps: acquiring an aircraft mixed structure fatigue life cumulative distribution function under a load spectrum according to a spectrum load fatigue life distribution of an aircraft mixed structure and each structural detail under the load spectrum; acquiring a CVM distribution distance between the aircraft mixed structure fatigue life cumulative distribution function under the load spectrum and a fatigue life distribution under a cutting spectrum, so that the distribution distance is within a set fatigue life distribution difference tolerance value; and selecting a maximum value in all low-load cutting levels corresponding to the distribution distance within the set fatigue life distribution difference tolerance value when low-load cutting is performed on an aircraft mixed structure fatigue test load spectrum. The application makes the mixed structure have the same fatigue life distribution under the cutting spectrum and the load spectrum.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of aircraft structure fatigue test, in particular to a low load truncation method and system for mixed structure fatigue test load spectrum. BACKGROUND

[0002] The aircraft structure fatigue test load spectrum contains a large number of low amplitude load cycles, and low load truncation as an important means of load spectrum simplification is of great significance to reduce the time and cost of the test.

[0003] In the prior art, for the durability fatigue test load spectrum considering only the fatigue crack initiation, the selection of the small load truncation level (SLTL) is mainly based on two ways: (1) taking the amplitude of SLTL as a certain percentage of the fatigue limit of the structural material; (2) taking the amplitude of SLTL as a certain percentage of the maximum load of the load spectrum, which is essentially consistent. The second way is an extension of the first way combined with the characteristics of the structural load. The fatigue failure behavior of mixed structures is more complex, and the research on the simplification of the load spectrum is relatively less. There is no reasonable solution to the problem of inconsistent SLTL for metal structures and composite structures or different material structures.

[0004] Therefore, there is a lack of a selection standard for the low load truncation level of mixed structures, so that the mixed structures have the same fatigue life distribution under the truncated spectrum and the load spectrum. SUMMARY

[0005] The embodiments of the present application provide a low load truncation method and system for mixed structure fatigue test load spectrum, so that the mixed structures have the same fatigue life distribution under the truncated spectrum and the load spectrum.

[0006] The embodiments of the present application provide a low load truncation method for mixed structure fatigue test load spectrum, comprising the following steps:

[0007] Performing fatigue test on the aircraft mixed structure and recording the load spectrum to obtain the spectrum load fatigue life distribution of each structural detail under the action of the load spectrum;

[0008] According to the spectrum load fatigue life distribution of the aircraft mixed structure and each structural detail under the action of the load spectrum, the fatigue life cumulative distribution function of the aircraft mixed structure under the action of the load spectrum is obtained;

[0009] determining a fatigue life distribution difference tolerance value for measuring the difference between the fatigue life cumulative distribution function of the aircraft hybrid structure under the action of the load spectrum and the fatigue life distribution under the action of the truncated spectrum; taking the CVM distribution distance between the fatigue life cumulative distribution function of the aircraft hybrid structure under the action of the load spectrum and the fatigue life distribution under the action of the truncated spectrum less than or equal to the fatigue life distribution difference tolerance value as the selection criterion; wherein the truncated spectrum is obtained by low-load truncation of the load spectrum at a low-load truncation level;

[0010] When low-load truncation is performed on the aircraft hybrid structure fatigue test load spectrum, the maximum value of all low-load truncation levels corresponding to the distribution distance meeting the selection criterion is selected.

[0011] Further, the aircraft hybrid structure is a series structure, a parallel structure or a series-parallel hybrid connection structure, and the determining step of the aircraft hybrid structure comprises:

[0012] determining the fatigue critical details of the hybrid structure according to experiments;

[0013] analyzing the logical relationship between the failure of the fatigue critical details and the failure of the overall structure to establish a hybrid structure failure logic model;

[0014] determining whether the aircraft hybrid structure is a series structure, a parallel structure or a series-parallel hybrid connection structure according to the hybrid structure failure logic model.

[0015] Further, the aircraft hybrid structure and the spectral fatigue life distribution of each structural detail under the action of the load spectrum are obtained to obtain the fatigue life cumulative distribution function of the aircraft hybrid structure under the action of the load spectrum, and the specific steps comprise:

[0016] The aircraft hybrid structure comprises k fatigue critical details, the first i fatigue critical detail is named D i , wherein ; under the action of the load spectrum, the fatigue life of the fatigue critical detail D i is , is a random variable and follows a lognormal distribution; the fatigue life distribution function of each fatigue critical detail is determined as , and the fatigue life of the aircraft hybrid structure is a function of , i.e. ;

[0017] When the aircraft hybrid structure is determined to be a series structure, the fatigue life under the action of the load spectrum is:

[0018] ;

[0019] The cumulative distribution function of the fatigue life of the aircraft hybrid structure is:

[0020] ;

[0021] When the aircraft hybrid structure is determined to be a parallel structure, the fatigue life under the load spectrum is:

[0022] ;

[0023] The cumulative distribution function of the fatigue life of the aircraft hybrid structure is:

[0024] ;

[0025] Wherein, is the fatigue life of the aircraft hybrid structure The cumulative distribution function of the fatigue life of the aircraft hybrid structure is:

[0026] Further, the CVM distribution distance between the fatigue life cumulative distribution function of the aircraft hybrid structure under the load spectrum and the fatigue life distribution under the truncated spectrum is less than or equal to the fatigue life distribution difference tolerance value as the selection criterion, and the specific steps include:

[0027] The fatigue life distribution of the aircraft hybrid structure under the load spectrum is F, and the fatigue life distribution under the truncated spectrum is G;

[0028] The distribution distance in the CVM distribution distance C Quantifies the difference between F and G;

[0029] ;

[0030] Wherein, is the fatigue life of the aircraft hybrid structure under the load spectrum, and are the cumulative distribution functions of F and G respectively, is the probability density function of F;

[0031] The distribution distance C Must be within the set fatigue life distribution difference tolerance value as the selection criterion:

[0032] ;

[0033] Wherein, is the fatigue life distribution difference tolerance value.

[0034] The embodiment of the application provides a low-load truncation system for a hybrid structure fatigue test load spectrum, which comprises:

[0035] A spectrum fatigue life distribution acquisition module for structural details is configured to perform fatigue tests on the mixed structure of the aircraft and record the load spectrum, and acquire the spectrum fatigue life distribution of each structural detail under the action of the load spectrum.

[0036] A mixed structure fatigue life cumulative distribution acquisition module is configured to acquire the fatigue life cumulative distribution function of the mixed structure of the aircraft under the action of the load spectrum according to the mixed structure of the aircraft and the spectrum fatigue life distribution of each structural detail under the action of the load spectrum.

[0037] A low-load truncation level determination module is configured to determine a fatigue life distribution difference tolerance value for measuring the difference between the fatigue life cumulative distribution function of the mixed structure of the aircraft under the action of the load spectrum and the fatigue life distribution under the action of the truncated spectrum, and take the CVM distribution distance between the fatigue life cumulative distribution function of the mixed structure of the aircraft under the action of the load spectrum and the fatigue life distribution under the action of the truncated spectrum less than or equal to the fatigue life distribution difference tolerance value as a selection criterion, wherein the truncated spectrum is obtained by performing low-load truncation on the load spectrum by selecting a low-load truncation level, and when performing low-load truncation on the fatigue test load spectrum of the mixed structure of the aircraft, the maximum value of all low-load truncation levels corresponding to the distribution distance meeting the selection criterion is selected.

[0038] The embodiment of the present application provides a low-load truncation method and system for a mixed structure fatigue test load spectrum, which has the following beneficial effects compared with the prior art:

[0039] According to the mixed structure of the aircraft and the spectrum fatigue life distribution of each structural detail under the action of the load spectrum, the fatigue life cumulative distribution function of the mixed structure of the aircraft under the action of the load spectrum is acquired, and the CVM distribution distance is used to measure whether the fatigue life cumulative distribution function of the mixed structure of the aircraft under the action of the load spectrum is equivalent to the fatigue life distribution under the action of the truncated spectrum, and when the distribution distance is within the set fatigue life distribution difference tolerance value, it is indicated that the fatigue life cumulative distribution function of the mixed structure of the aircraft under the action of the load spectrum is equivalent to the fatigue life distribution under the action of the truncated spectrum, and the maximum value of all low-load truncation levels corresponding to this time is taken as the selection standard of low-load truncation. BRIEF DESCRIPTION OF DRAWINGS

[0040] Figure 1 A flowchart of the low-load truncation method for the mixed structure fatigue test load spectrum is provided for the present application.

[0041] Figure 2 A mixed structure fatigue life distribution calculation flow provided for the embodiment of the present application is provided.

[0042] Figure 3 A mixed structure connection structure model of multiple fatigue dangerous details is provided for the embodiment of the present application.

[0043] Figure 4The material fatigue performance and fatigue risk detail stress spectrum distribution characteristics provided for the embodiment of the present application;

[0044] Figure 5 The specimen layout provided for the embodiment of the present application;

[0045] Figure 6 The connecting plate component material P-S-N curve provided for the embodiment of the present application;

[0046] Figure 7 The relationship diagram between the flight overload, the structure load and the fatigue risk detail stress response provided for the embodiment of the present application;

[0047] Figure 8 The spectrum fatigue life distribution of the fatigue risk detail and the structure provided for the embodiment of the present application;

[0048] Figure 9 The influence of the low load truncation on the fatigue life distribution provided for the embodiment of the present application;

[0049] Figure 10 The test load spectrum provided for the embodiment of the present application;

[0050] Figure 11 The diagram of the spectrum load fatigue life distribution changing with the SLTL provided for the embodiment of the present application;

[0051] Figure 12 The low load truncation method specific implementation step flow chart of a mixed structure fatigue test load spectrum provided for the embodiment of the present application. DETAILED DESCRIPTION

[0052] In order to make the above objectives, features and advantages of the present application more apparent and comprehensible, the following will combine the attached drawings for the specific implementation manner of the present application to make a detailed description. Figures 1-12 The specific implementation manner of the present application is described in detail. In the following description, a lot of specific details are set forth in order to facilitate a full understanding of the present application. However, the present application can be implemented in many other ways different from the description herein, and the person skilled in the art can make similar improvements without departing from the connotation of the present application, therefore the present application is not limited by the specific embodiments disclosed below.

[0053] As shown in Figure 1 , the low load truncation method of a mixed structure fatigue test load spectrum provided by the embodiment of the present application comprises the following steps:

[0054] Step one: fatigue test is performed on the mixed structure of the airplane and the load spectrum is recorded, and the spectrum load fatigue life distribution of each structure detail under the action of the load spectrum is obtained.

[0055] Step two: Obtain the fatigue life cumulative distribution function of the aircraft hybrid structure under the load spectrum according to the spectrum fatigue life distribution of each structure detail under the load spectrum.

[0056] Step three: Determine the fatigue life distribution difference tolerance value for measuring the difference between the fatigue life cumulative distribution function of the aircraft hybrid structure under the load spectrum and the fatigue life distribution under the truncated spectrum; take the CVM (Cramer-Von Mises) distribution distance between the fatigue life cumulative distribution function of the aircraft hybrid structure under the load spectrum and the fatigue life distribution under the truncated spectrum less than or equal to the fatigue life distribution difference tolerance value as the selection criterion; wherein the truncated spectrum is obtained by low-load truncation of the load spectrum by selecting a low-load truncation level; when the low-load truncation is performed on the aircraft hybrid structure fatigue test load spectrum, select the maximum value of all low-load truncation levels corresponding to the distribution distance that meets the selection criterion.

[0057] I. Low-load truncation level selection criterion

[0058] Considering the randomness of structural fatigue life, the load spectrum simplification should not change the fatigue life distribution characteristics under the original load spectrum. When the fatigue life distribution of the aircraft hybrid structure under the load spectrum is F, and the fatigue life distribution under the truncated spectrum is G, it can be considered that the simplified load spectrum and the truncated spectrum are equivalent when F and G have approximately the same distribution. Referring to the Cramer-von Mises statistic, the distribution distance C is used to evaluate the difference between F and G:

[0059] ;

[0060] In the formula, is the fatigue life of the aircraft hybrid structure under the load spectrum, and are the cumulative distribution functions of distributions F and G, respectively, is the probability density function of distribution F, and the distribution distance C can be solved by numerical integration method after determining the parameters of each distribution.

[0061] Distribution distance C needs to be within the set fatigue life distribution difference tolerance value as the selection criterion, wherein is a given small quantity greater than 0, and if the following formula is true, it is considered that F and G approximately obey the same distribution:

[0062] ;

[0063] In the formula, is the tolerance value of the distribution function difference, which can generally be taken as .

[0064] By calculating the fatigue life distribution under different SLTL, the change rule of distribution distance C with SLTL can be established, generally C increases with the increase of SLTL, and gradually tends to the limit 1 / 3. The selection of SLTL not only follows the principle of accuracy of fatigue life evaluation, but also follows the principle of test economy, and needs to shorten the length of the simplified spectrum as much as possible, and the length of the simplified spectrum shortens with the increase of SLTL, so the critical SLTL level of can be taken.

[0065] Assuming that the fatigue performance evaluation preliminarily determines that the mixed structure of the aircraft contains k fatigue critical details, the first i fatigue critical detail is named D i , wherein . Through load response analysis by test monitoring, simulation calculation or theoretical analysis, the relationship between aircraft overload, local structure load and detail stress response can be established, when the aircraft overload is n , the local structure load is , and the stress response of the fatigue critical detail D i is , then the local structure load spectrum and the fatigue critical detail stress response spectrum can be obtained according to the flight overload spectrum of the aircraft. Since the structure load spectrum has randomness, correspondingly, the stress response spectrum of each fatigue critical detail in the structure also has randomness, in addition, the fatigue performance of each fatigue critical detail has dispersion, which leads to the dispersion of the fatigue life of each fatigue critical detail, and finally the fatigue life of the mixed structure also shows significant dispersion. Assuming that each fatigue critical detail is independent of each other, under the original load spectrum, the fatigue life of the fatigue critical detail D i is , is a random variable, generally considered to follow a lognormal distribution, and its distribution function is , the fatigue life of the mixed structure of the aircraft is a function of , that is .

[0066] When the SLTL of the overload spectrum is , the SLTL of the corresponding structure load spectrum is , and the SLTL of the corresponding fatigue critical detail stress response spectrum is . Due to the difference in stress response and fatigue performance between different fatigue critical details, the low load cut-off has different effects on the fatigue life distribution of each fatigue critical detail, and the fatigue critical detail D i The distribution distance of the fatigue life distribution under the cut-off spectrum and the original spectrum is Ci If there are also strict requirements for the evaluation and verification of fatigue performance of fatigue hazard details, then it is also necessary to establish SLTL selection criteria for fatigue hazard details. Among them, the tolerance value for the difference in fatigue life distribution in terms of fatigue risk details. Given by the design, we generally only focus on the impact of low load cutoff on the fatigue life distribution of the structure, that is, we use the formula as the selection criterion for SLTL of mixed structures with multiple fatigue risk details.

[0067] II. Calculation Method for Spectral Loaded Fatigue Life Distribution

[0068] 1. Fatigue life distribution calculation model for aircraft hybrid structures with multiple fatigue risk details

[0069] like Figure 2 The diagram shows the calculation process for the fatigue life distribution of a hybrid aircraft structure with multiple fatigue risk details. The dispersion of the fatigue life of the hybrid aircraft structure originates from the dispersion of the fatigue life of the fatigue risk details. Based on the impact of fatigue risk detail failure on the overall structural failure behavior, or the connection relationship between various fatigue risk details, the hybrid aircraft structure can be simplified into a series structure model, a parallel structure model, or a series-parallel hybrid connection structure model, as shown below. Figure 3 As shown.

[0070] The specific steps for determining whether an aircraft hybrid structure is a series structure, a parallel structure, or a hybrid series-parallel connection structure include: identifying fatigue-prone details of the aircraft hybrid structure based on experiments; analyzing the logical relationship between the failure of fatigue-prone details and the failure of the overall structure, and establishing a failure logic model for the aircraft hybrid structure; the failure logic model of the aircraft hybrid structure can characterize whether the aircraft hybrid structure is a series structure, a parallel structure, or a hybrid series-parallel connection structure. If the failure of any fatigue-prone detail leads to the failure of the entire structure, then the aircraft hybrid structure is a series structure. If any fatigue-prone detail in the aircraft hybrid structure has load-bearing capacity, and when one fatigue-prone detail fails, the other fatigue-prone details in the aircraft hybrid structure continue to bear load until the last remaining fatigue-prone detail fails, then the aircraft hybrid structure is a parallel structure. When both of the above two failure logics from fatigue-prone details to the aircraft hybrid structure exist, the aircraft hybrid structure is a hybrid series-parallel structure.

[0071] Aircraft hybrid structure includes k The first detail of fatigue risk will be the first i A detail of fatigue risk was named D i ,in Fatigue hazard details under load spectrum D i The fatigue life is , Let the fatigue life distribution function of each fatigue critical detail be , the fatigue life of the hybrid structure is a function of , i.e. .

[0072] For the series structure model, the fatigue life of the hybrid structure of the aircraft is , then the cumulative distribution function of is:

[0073] ;

[0074] For the parallel structure model, if the load redistribution process after the failure of a fatigue critical detail is not considered, the fatigue life of the hybrid structure of the aircraft is , then the cumulative distribution function of is:

[0075] ;

[0076] The calculation of the fatigue life distribution of the parallel structure considering the load redistribution process is more complex, and there is no explicit analytical solution, which is generally calculated by numerical simulation method. Similarly, for the series-parallel hybrid connection structure, as long as the fatigue life distribution of each fatigue critical detail is determined, the fatigue life distribution of the hybrid structure of the aircraft can also be obtained. When the hybrid structure of the aircraft is determined to be a series-parallel hybrid connection structure, the spectral fatigue life distribution function of the hybrid structure is established according to the failure logic model of the hybrid structure.

[0077] Therefore, according to the response of the structure details calculated by the full aircraft load, combined with the fatigue performance of the material of the fatigue critical details, the fatigue life distribution of the fatigue critical details can be calculated, and then according to the fatigue failure logic of the fatigue critical details to the structure, the fatigue life distribution of the hybrid structure can be established.

[0078] 2. Sampling calculation method of spectral fatigue life distribution of fatigue critical details

[0079] The dispersion of the fatigue life of the fatigue critical details mainly comes from the dispersion of the material fatigue performance and the detail stress spectrum, and the fatigue life distribution is difficult to be solved directly by analytical method. Based on the distribution characteristics of the material fatigue performance and the stress spectrum of the fatigue critical details, the spectral fatigue life distribution can be estimated by Monte Carlo sampling method.

[0080] As shown in Fig. (a) in Figure 4 , the dispersion of the material fatigue performance can be described by the probability-stress-fatigue life curve (P-S-N curve), and the fatigue life of the material at each stress level obeys the lognormal distribution, i.e. , the conditional fatigue limit (corresponding to the cycle base of the fatigue strength) obeys normal distribution, i.e. , while the S-N curve p corresponding to the probability obeys Weibull distribution, i.e. p The conditional fatigue limit based on the cumulative probability p and the fatigue life data under each level of load based on the cumulative probability are fitted by least square method.

[0081] Figure 4 As shown in Fig. (b) in , the dispersion of the stress spectrum can be described by the probability-stress-exceedance frequency curve (P-S-E curve), the exceedance frequency of the stress spectrum under each level of stress obeys lognormal distribution, i.e. q , while the stress-exceedance frequency curve corresponding to the probability q is fitted based on the exceedance frequency data under each level of stress.

[0082] According to Miner theory, the fatigue life D i of a fatigue critical detail under a certain load spectrum can be calculated by the following formula:

[0083]

[0084] wherein, is the fatigue critical detail D i under constant amplitude stress cycle , is the stress amplitude, is the stress mean value, is the probability density function of the load , and is the integral interval, representing the load interval in the load spectrum that needs to be analyzed for damage. When the mean stress is constant, ( Q ) is the probability density function of the stress amplitude. When the stress-exceedance frequency curve is known, there is

[0085] , and the spectrum fatigue life can be expressed as: l S

[0086]

[0087] wherein, the integral interval ~ is the stress amplitude interval in the stress spectrum that needs to be analyzed for fatigue damage, and for the original spectrum has​​​​​​​ and for low load spectrum , is the maximum stress amplitude in the stress spectrum.

[0088] According to the formula, as long as the material fatigue performance curve equation is given based on the sampling of the distribution characteristics of the material fatigue performance and the load spectrum and the load-exceeding frequency curve equation , the corresponding sampling results of the spectrum load fatigue life can be calculated. Assuming obeys the lognormal distribution (base 10), according to a large number of sampling results, the maximum likelihood estimation method can be used to estimate the corresponding distribution parameters.

[0089] III. Numerical examples and experimental verification

[0090] 1. Structure and material

[0091] 7050 aluminum alloy / 1045 steel hybrid single-pin single-lap connectors are used for low load spectrum cutting test research of multi-detail hybrid material connection structure, the sample layout is shown in Figure 5 , the bolt material is 30CrMnSiA ultra-high strength steel. The sample is mainly subjected to tensile fatigue load, without considering other unconventional failure modes, the preliminary assessment of the potential fatigue risk site is the bolt hole A of the aluminum alloy plate and the bolt hole B of the steel plate, the failure of any risk site will lead to the failure of the entire structure, so the sample can be regarded as a series structure of two fatigue risk details, the aluminum alloy plate and the steel plate are defined as fatigue risk details D 1 and D 2 respectively.

[0092] The material composition and basic mechanical properties of the connecting plate are shown in Table 1. The P-S-N curve of the fatigue performance of the connecting plate material is established through the test data (stress ratio K t =-1) of the smooth sample (stress concentration coefficient R =-1), as shown in Fig. (a) in Figure 6 , the P-S-N curve of the fatigue performance of 7050 aluminum alloy, and as shown in Fig. (b) in Figure 6 , the P-S-N curve of the fatigue performance of 1045 steel.

[0093] Table 1 Material composition and basic mechanical properties of connector components

[0094]

[0095] 2. Load spectrum

[0096] The original load spectrum was obtained by processing the measured center-of-gravity normal overload spectrum data of the Y-8 transport aircraft. Each spectrum represents 5000 flights, and the average overload during the flight is [data missing]. n z,m =1g, overload amplitude n z,a The probability-overload-over frequency curve is as follows: Figure 7 Figure (a) shows the probability-load-overfrequency curve of the center-of-gravity normal overload spectrum under 5000 flights. The minimum overload amplitude in the original spectrum is taken as 0.1g, and the total number of cycles in the median spectrum is 1091917.

[0097] Based on the relationship between flight loads and the stress response of structural details, the load spectrum of the specimen can be further designed. Assuming a key structural component of the aircraft uses the method shown in Figure [Figure Number Missing]... Figure 5 The structure shown is designed to address fatigue-prone stress levels. S With external load n z Linear relationship ,in For details regarding the fatigue risk of the structure, the nominal stress under 1g flight conditions, and the nominal stress of the aluminum alloy plate under 1g flight conditions. =50MPa, the nominal stress of the steel plate is =125MPa, from which the relationship between flight overload, structural load, and stress response in fatigue-prone details can be established, such as Figure 7 Figure (b) shows the stress response relationship between flight overload, structural load, and fatigue hazard details, where the sample load corresponding to a 1g overload is... P =3.3kN. Combined with Figure 7 Figures (a) and (b) in the figure can transform the aircraft overload spectrum into the specimen load spectrum (probability-load-exceedance frequency curve) and the stress spectrum (probability-stress-exceedance frequency curve) of the fatigue hazard details, which can be used for test load spectrum design and fatigue life calculation of fatigue hazard details, respectively.

[0098] 3. Calculation of fatigue life distribution in the original spectrum

[0099] Fatigue performance analysis of connection structures is generally performed based on the stress severity factor method (SSF method). After calculating the SSF value of each bolt hole edge, the SSF value is regarded as the theoretical stress concentration factor. K t Combining with the same K tThe simple notched P-S-N curve of the fatigue critical detail and the fatigue dangerous detail stress spectrum P-S-E curve, based on the fatigue dangerous detail stress spectrum fatigue life distribution sampling calculation method, can be used to calculate the spectrum fatigue life distribution of each fatigue dangerous detail, and finally the spectrum fatigue life distribution of the hybrid material connection structure specimen can be calculated according to the fatigue life distribution model of the series structure.

[0100] It is assumed that the fatigue life of the fatigue critical detail D 1 and the fatigue life of the fatigue dangerous detail D 2 both obey the lognormal distribution, and are F 1 and F 2 respectively. According to the formula, the cumulative distribution function of the fatigue life of the connection structure is:

[0101]

[0102] The stress severity factor of hole A is 5.6 and the stress severity factor of hole B is 4.3, which are calculated based on the SSF method. By using the Peterson fatigue notch coefficient correction method and the Goodman average stress correction method, the P-S-N curve of the material in Figure 6 can be corrected to obtain the P-S-N curve of each detail material under the specified average stress and the specified SSF value. Based on the material P-S-N curve and the stress spectrum P-S-E curve, the fatigue life distribution of each fatigue critical detail is established by a large number of sampling calculations, and the sampling number is taken as 1000. According to the sampling calculation results, the fatigue life distribution of each fatigue critical detail is fitted, and the fatigue life distribution of the hybrid connection structure is calculated according to the series structure fatigue life distribution calculation model. The calculation results are shown in Figure 8 , which is the spectrum fatigue life distribution of the fatigue critical detail and the structure. It can be seen that the fatigue life distribution of the hybrid structure is still close to the lognormal distribution, and is similar to the fatigue life distribution of the fatigue critical detail D 1. At this time, the probability of the fatigue critical detail D 1 failing first is 79%.

[0103] 4. Analysis of low load cut-off level selection

[0104] The spectrum fatigue life distribution of each fatigue critical detail and the hybrid structure under different is calculated, and the change of the fatigue life distribution with is shown in Fig. (a) in Figure 9 . Further, the distribution distance C , C of the fatigue life distribution under the cut-off spectrum relative to the original spectrum fatigue life distribution is calculated, and the change trend of the distribution distance with is shown in Fig. (b) in Figure 9 . It can be seen that in ​At lower loads, low-load cutoff has no effect on the fatigue life distribution of the hybrid structure. As the SLTL increases, the fatigue life of the hybrid structure gradually increases, and the distribution distance... C It also gradually increased.

[0105] Make details D 2. Significant changes in lifetime distribution It must be higher than the details that make fatigue dangerous. D 1. Significant changes in lifetime distribution The median value of the material conditional fatigue limit (corresponding to the cycle base) obtained by means of stress correction and fatigue notch factor correction. ) are respectively =17.9MPa, =71.6MPa, the external load required for the nominal stress response of the fatigue fault detail to reach the median value of the conditional fatigue limit is defined as the fatigue limit load (FLL). Calculate the details of fatigue hazard separately. D 1 and D 2's FLL is =1.18kN, =1.89kN, such as Figure 9 As shown in Figure (a), the life distribution of each fatigue-prone detail in the structure reaches the corresponding... Significant changes only gradually began to occur later. Therefore, under the same cutoff spectrum, i.e., the same... Below, the degree of deviation of the fatigue life distribution of each fatigue risk detail from the original spectrum fatigue life distribution is... Relevant, generally speaking The lower the level of fatigue risk detail, the greater the deviation in fatigue life distribution. Figure 9 As can be seen from Figure (b), for the same The difference in fatigue life distribution between the truncated and original spectra for the hybrid structure is slightly smaller than that for the detailed spectra. D 1. The distribution of fatigue life varies, and is far greater than the details of fatigue hazard. D 2. Differences in fatigue life.

[0106] Assuming the structural design requirements focus only on the fatigue performance assessment of the hybrid structure, the influence of the cutoff spectrum on the fatigue life distribution of fatigue risk details can be ignored, and the distribution distance can be initially taken as... C tolerance value Thus, it was determined =1.35kN, from Figure 9 Figure (a) shows that Compare Slightly higher, in this The fatigue life distribution of the detail will not change significantly. According to Figure 7 Figure (b) in the figure, the SLTL of the corresponding overload spectrum is =0.41g, the SLTL of the fatigue detail stress spectrum is =20.6MPa, =51.5MPa, that is, the SLTL corresponds to 104% of the fatigue limit of aluminum alloy 70% of the fatigue limit of steel According to Figure (a) in the figure Figure 7 The total cycle distribution of the corresponding truncated spectrum is , wherein =4.14, =0.215, the cycle number of the median load spectrum is reduced from 1.09 million to 1.38 million, and the total cycle number is 1.26% of the original, which will greatly save the time and cost of fatigue test.

[0107] 5. Test verification

[0108] The fatigue test under the original spectrum and the truncated spectrum is designed. The original spectrum is represented as L0, and the analysis is selected according to the low load truncation level Select =1.35kN, the truncated spectrum obtained is represented as L1, and The corresponding =2.25kN, the truncated spectrum obtained is represented as L2, and the details of each load spectrum are shown in Table 2, which is the cycle number of the load spectrum under different SLTL.

[0109] Table 2 Cycle number of load spectrum under different SLTL

[0110]

[0111] The test of 8 samples under each type of load spectrum is completed. In order to consider the variability of the load, based on the load spectrum distribution given in Figure (a) in the figure Figure 7 , 8 load spectrum curves are randomly sampled, as shown in Figure (a) in the figure Figure 10 , which is the test load spectrum design (load mean Pm=3.3kN), based on which the original spectrum L0 and the truncated spectrum L1, L2 are designed, and the random load spectrum is generated for the loading of each sample, as shown in Figure (b) in the figure Figure 10 , which is a test load spectrum segment (L0), the test is completed on the MTS 370 fatigue test system, and the test frequency is set to 10Hz.

[0112] The experimental results are shown in Table 3, which presents the fatigue test results under different spectra. Those marked with "*" indicate fatigue failure at the edge of the hole in the steel plate, while the rest indicate fatigue failure at the edge of the hole in the aluminum alloy plate. The maximum likelihood estimation method was used to provide different... The fatigue life distribution of the hybrid structure, the spectral fatigue life distribution varies with Changes such as Figure 11 As shown in Table 4, the distribution distances of fatigue life distributions under the cutoff spectra L1 and L2 and the original spectrum L0 were examined. This serves as a test of the simplification effect of the cutoff spectra. It can be seen that the distribution distances obtained experimentally under each cutoff spectrum are basically on the same order of magnitude as the theoretical distribution distances, with small differences. This indicates the effectiveness of using distribution distance to evaluate the impact of low-load cutoff on fatigue life distribution. The fatigue life distributions under L1 and L0 approximately conform to the same distribution, while the fatigue life distributions under L2 and L0 differ significantly. Therefore, the tolerance value of the fatigue life distribution distance is... This was to determine that SLTL can achieve good load spectrum simplification.

[0113] Table 3 Fatigue test results under different spectra

[0114]

[0115] Table 4. Verification of the spectral simplification effect of truncation

[0116]

[0117] 6. Discussion

[0118] The probability of first failure in the aircraft's hybrid structure W f The most significant fatigue hazard detail is defined as the primary hazard detail (PHD). Clearly, for tandem structures, the fatigue life distribution of the aircraft hybrid structure is closest to that of the PHD. Furthermore, the higher the probability of the PHD's initial failure, the greater the similarity between the hybrid structure's fatigue life distribution and the PHD's fatigue life distribution. This implies that other fatigue hazard details have a smaller impact on the structural fatigue life distribution.

[0119] For the 7050 aluminum alloy / 1045 steel hybrid single-nail single-lap connector of this invention, fatigue life distribution calculations based on the original spectrum reveal fatigue hazard details. D 1 probability of failure first W f =79%, details of fatigue hazards D 2's probability of early failure W f =21%, then PHD represents fatigue hazard details. D1. Assuming the original load spectrum and other structural dimensions remain unchanged, only the thickness of the steel plates in the hybrid structure is varied. Analyze the fatigue hazard details and fatigue life distribution of the hybrid structure under different steel plate thicknesses. The calculation range for steel plate thickness is 1.1mm~1.5mm. Calculate the fatigue hazard details under different steel plate thicknesses. D 1 probability of early failure W f Calculation results show that the PHD of the hybrid structure under different steel plate thicknesses is a fatigue-prone detail. D 1. Further calculation of fatigue risk details under different steel plate thicknesses. D The distribution distance C between the fatigue life distribution of 1 and the fatigue life distribution of the hybrid structure was determined, and fatigue hazard details were established. D 1. Probability of failure first W f The relationship between the distribution distance C and the probability of PHD's early failure. W f When >85%, the distribution distance C < 10. -3 In this case, the fatigue life distribution of the PHD can be considered to be the same as that of the hybrid structure, and the influence of other details on the fatigue life distribution can be ignored. Correspondingly, even if the fatigue life distribution of other minor fatigue-prone details changes greatly under the low-load cutoff spectrum, the influence on the fatigue life distribution of the hybrid structure is very limited. At this time, the load spectrum simplification problem of the hybrid structure with multiple fatigue-prone details can be approximated as the traditional single-detail load spectrum simplification problem.

[0120] IV. Conclusion

[0121] 1) Considering the fatigue performance of materials and the dispersion of load spectrum, a fatigue life distribution calculation model for a hybrid structure with multiple fatigue risk details was established. The distribution distance was used to evaluate the impact of low load removal on the fatigue life distribution of the hybrid structure. The corresponding low load removal level of the hybrid structure can be determined according to the tolerance value given in the design.

[0122] 2) Test results of aluminum alloy-steel hybrid single-nail connection structures show that, based on the tolerance value for fatigue life distribution differences... The determined structural low-load cutoff level The fatigue limit load of PHD, which is the primary fatigue hazard detail. The load spectrum is slightly higher, but the length is greatly shortened without significantly changing the fatigue life distribution of the hybrid structure. The simplification effect is satisfactory, and this method can be applied to the simplification of the fatigue test load spectrum of large-size complex aircraft structures.

[0123] 3) Under the same cutoff spectrum, fatigue limit load The fatigue life distribution of lower fatigue hazard details is significantly affected. This is especially true when the probability of the primary fatigue hazard detail (PHD) failing first in a hybrid structure is lower.W f When greater than 85%, the fatigue life of the hybrid structure is approximately identically distributed with the fatigue life of the PHD, and the influence of the remaining minor fatigue hazard details on the fatigue life distribution of the hybrid structure can be ignored, and the simplification problem of the hybrid structure load spectrum can be approximately regarded as the simplification problem of the traditional single-detail load spectrum.

[0124] The embodiment of the application provides a low-load truncation system for a hybrid structure fatigue test load spectrum, comprising:

[0125] A spectrum fatigue life distribution acquisition module for a structure detail is configured to perform a fatigue test on an aircraft hybrid structure and record a load spectrum, and acquire a spectrum fatigue life distribution of each structure detail under the action of the load spectrum.

[0126] A hybrid structure fatigue life cumulative distribution acquisition module is configured to acquire a fatigue life cumulative distribution function of the aircraft hybrid structure under the action of the load spectrum according to the spectrum fatigue life distribution of each structure detail under the action of the load spectrum.

[0127] A low-load truncation level determination module is configured to determine a fatigue life distribution difference tolerance value for measuring the difference between the fatigue life cumulative distribution function of the aircraft hybrid structure under the action of the load spectrum and the fatigue life distribution under the action of a truncated spectrum, and determine a CVM distribution distance between the fatigue life cumulative distribution function of the aircraft hybrid structure under the action of the load spectrum and the fatigue life distribution under the action of the truncated spectrum as a selection criterion, wherein the truncated spectrum is obtained by performing low-load truncation on the load spectrum at a low-load truncation level, and when performing low-load truncation on the aircraft hybrid structure fatigue test load spectrum, the maximum value of all low-load truncation levels corresponding to the distribution distance meeting the selection criterion is selected.

[0128] One specific implementation is as follows:

[0129] As shown in Figure 12 , the embodiment discloses a low-load truncation method for a hybrid structure fatigue test load spectrum, and the specific steps are as follows:

[0130] S1, acquire the probability-stress-fatigue life P-S-N curve and the probability-stress-exceeding frequency P-S-E curve of each structure detail.

[0131] S2, use Monte Carlo sampling on the P-S-N curve and the P-S-E curve respectively, use the linear cumulative damage theory Miner to acquire the spectrum fatigue life of each structure detail under the action of the load spectrum according to the sampling result, and use the logarithmic normal distribution to fit the spectrum fatigue life distribution of each structure detail under the action of the load spectrum.

[0132] S3, determining the aircraft hybrid structure as a series structure, a parallel structure or a series-parallel hybrid connection structure.

[0133] S4, obtaining an aircraft hybrid structure fatigue life cumulative distribution function under a load spectrum according to the determined aircraft hybrid structure and the spectrum fatigue life distribution of each structure detail under the load spectrum.

[0134] S5, selecting a low load cut-off level to cut off the low load of the load spectrum to obtain a cut-off spectrum; and quantifying the difference between the aircraft hybrid structure fatigue life cumulative distribution function under the load spectrum and the fatigue life distribution under the cut-off spectrum by CVM distribution distance, so that the quantification result is within the set fatigue life distribution difference tolerance value.

[0135] S6, when the low load of the aircraft hybrid structure fatigue test load spectrum is cut off, selecting the low load cut-off level corresponding to the quantification result within the set fatigue life distribution difference tolerance value.

[0136] The above embodiments only express several embodiments of the present application, and the description is more specific and detailed, but it cannot be understood as a limitation on the scope of the patent. It should be noted that for ordinary skilled persons in the art, without departing from the concept of the present application, several modifications and improvements can be made, which belong to the protection scope of the present application. Therefore, the protection scope of the patent of the present application should be subject to the appended claims.

Claims

1. A low-load truncation method for hybrid spectrum fatigue test load spectrum, characterized in that, The method comprises the following steps: fatigue test is performed on the aircraft hybrid structure, and a load spectrum is recorded to obtain a spectrum load fatigue life distribution of each structural detail under the action of the load spectrum; an aircraft hybrid structure fatigue life cumulative distribution function under the action of the load spectrum is obtained according to the aircraft hybrid structure and the spectrum load fatigue life distribution of each structural detail under the action of the load spectrum; a fatigue life distribution difference tolerance value for measuring the difference between the aircraft hybrid structure fatigue life cumulative distribution function under the action of the load spectrum and the fatigue life distribution under the action of the truncated spectrum is determined; a CVM distribution distance between the aircraft hybrid structure fatigue life cumulative distribution function under the action of the load spectrum and the fatigue life distribution under the action of the truncated spectrum is less than or equal to the fatigue life distribution difference tolerance value as a selection criterion; wherein the truncated spectrum is obtained by low-load truncation of the load spectrum by selecting a low-load truncation level; when the low-load truncation of the aircraft hybrid structure fatigue test load spectrum is performed, the maximum value of all low-load truncation levels corresponding to the distribution distance meeting the selection criterion is selected.

2. The low-load truncation method of hybrid structural fatigue test load spectrum of claim 1, wherein, The aircraft hybrid structure is a series structure, a parallel structure or a series-parallel hybrid connection structure, and the determination step comprises: determining a fatigue critical detail of the hybrid structure according to an experiment; analyzing a logical relationship between failure of the fatigue critical detail and failure of the overall structure to establish a hybrid structure failure logic model; determining that the aircraft hybrid structure is a series structure, a parallel structure or a series-parallel hybrid connection structure according to the hybrid structure failure logic model.

3. The low-cycle cut-off method of hybrid structural fatigue test load spectrum according to claim 1, characterized in that, The specific steps of obtaining the aircraft hybrid structure fatigue life cumulative distribution function under the action of the load spectrum according to the aircraft hybrid structure and the spectrum load fatigue life distribution of each structural detail under the action of the load spectrum comprise: The aircraft hybrid structure contains k fatigue details, the first i fatigue detail is named D i , wherein ; under the load spectrum, the fatigue detail D i has a fatigue life of , is a random variable and obeys a lognormal distribution; the fatigue life distribution function of each fatigue detail is determined as , and the fatigue life of the aircraft hybrid structure is a function of , that is ; when it is determined that the aircraft hybrid structure is a series structure, the fatigue life under the action of the load spectrum is: ； The cumulative distribution function of X is: ； when it is determined that the aircraft hybrid structure is a parallel structure, the fatigue life under the action of the load spectrum is: ； The cumulative distribution function of X is: ； wherein Fatigue life of aircraft hybrid structures Cumulative distribution function.

4. The low-cycle cut-off method of hybrid structural fatigue test load spectrum according to claim 1, characterized in that, The specific steps of taking the CVM distribution distance between the aircraft hybrid structure fatigue life cumulative distribution function under the action of the load spectrum and the fatigue life distribution under the action of the truncated spectrum less than or equal to the fatigue life distribution difference tolerance value as the selection criterion comprise: determining that the fatigue life distribution of the aircraft hybrid structure under the action of the load spectrum is F, and the fatigue life distribution under the action of the truncated spectrum is G; Using CVM distribution distance in distribution distance C Quantifying the difference between F and G; ； wherein, is the fatigue life of the aircraft hybrid structure under the load spectrum, and are the cumulative distribution functions of F and G, respectively, is the probability density function of F. Distribution distance C Need to be within the set fatigue life distribution difference tolerance value as selection criteria: ； wherein, is the fatigue life distribution difference tolerance value.

5. A low-load truncation system for hybrid spectrum fatigue test load spectrum, characterized in that, The method comprises the following steps: a spectrum load fatigue life distribution obtaining module of a structural detail, configured to perform fatigue test on the aircraft hybrid structure, record a load spectrum, and obtain a spectrum load fatigue life distribution of each structural detail under the action of the load spectrum; an aircraft hybrid structure fatigue life cumulative distribution obtaining module, configured to obtain an aircraft hybrid structure fatigue life cumulative distribution function under the action of the load spectrum according to the aircraft hybrid structure and the spectrum load fatigue life distribution of each structural detail under the action of the load spectrum; The low-load truncation level determination module is configured to determine a fatigue life distribution difference tolerance value for measuring the difference between the fatigue life cumulative distribution function of the aircraft hybrid structure under the load spectrum and the fatigue life distribution under the truncated spectrum; and determine a selection criterion that the CVM distribution distance between the fatigue life cumulative distribution function of the aircraft hybrid structure under the load spectrum and the fatigue life distribution under the truncated spectrum is less than or equal to the fatigue life distribution difference tolerance value; wherein the truncated spectrum is obtained by performing low-load truncation on the load spectrum by selecting a low-load truncation level; and when performing low-load truncation on the aircraft hybrid structure fatigue test load spectrum, selecting the maximum value of all low-load truncation levels corresponding to the distribution distance that meets the selection criterion.

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