An aero-part damage low-cycle fatigue life evaluation method and system

Abstract

A method and system for assessing the low-cycle fatigue life of aerospace components, relating to the field of component life assessment technology, includes: acquiring the current damage depth and working stress of the aerospace component to be assessed; acquiring a set of S-N curves for the same material as the aerospace component to be assessed, including damage S-N curves at at least two different damage depths; calculating the crack initiation life by interpolating the set of S-N curves based on the current damage depth and working stress; using the sum of the current damage depth and a preset correction value as the initial crack size, performing crack propagation analysis based on the initial crack size to obtain the initial crack propagation life, and conservatively processing the initial crack propagation life to obtain the target crack propagation life; and summing the crack initiation life and the target crack propagation life to obtain the total low-cycle fatigue life of the aerospace component to be assessed. Implementing this technical solution improves the reliability of life assessment.

Description

Technical Field

[0001] This application relates to the field of component life assessment technology, specifically to a method and system for assessing the low-cycle fatigue life of aerospace components. Background Technology

[0002] During aircraft service, various aerospace components (aircraft fuselage, landing gear, engines) are susceptible to adverse factors such as impacts from external objects and assembly collisions, resulting in surface scratches, abrasions, and nicks, which affect the strength of the components. Currently, the assessment of the remaining service life of damaged components mainly uses the crack propagation method. This method has two problems. First, since the permissible size of service damage is generally small (e.g., depth < 0.38 mm), directly equating it to an initial crack for calculation is affected by the "small crack effect," leading to a significant deviation between theoretical calculation results and actual physical behavior, resulting in insufficient reliability. Second, in the actual low-cycle fatigue process of damage, there is a crack initiation process. The crack propagation method ignores this process, leading to overly conservative life calculation results and potentially unnecessary component replacements. Therefore, the low-cycle fatigue life assessment methods for aerospace component damage in related technologies suffer from low reliability. Summary of the Invention

[0003] To address the aforementioned technical problems, this application provides a method and system for assessing the low-cycle fatigue life of aircraft parts.

[0004] In a first aspect, this application provides a method for assessing the low-cycle fatigue life of damaged aerospace components, comprising: acquiring the current damage depth and working stress of the aerospace component to be assessed; acquiring a set of SN curves of a sample material of the same material as the aerospace component to be assessed, wherein the set of SN curves includes damage SN curves of the sample material at at least two different damage depths; calculating the crack initiation life by interpolation of the set of SN curves based on the current damage depth and working stress; using the sum of the current damage depth and a preset correction value as the initial crack size, performing crack propagation analysis based on the initial crack size to obtain the initial crack propagation life, and performing a conservative treatment on the initial crack propagation life to obtain the target crack propagation life, wherein the conservative treatment is used to represent the reduction treatment of the initial crack propagation life; and summing the crack initiation life and the target crack propagation life to obtain the total low-cycle fatigue life of the aerospace component to be assessed.

[0005] By employing the above technical solutions, the current damage depth and working stress of the aerospace component to be evaluated, as well as the damage SN curves of sample materials of the same material at at least two different damage depths, can be obtained, providing basic data for subsequent life assessment. Interpolation calculations of a set of SN curves yield the crack initiation life, taking into account the crack initiation process and avoiding overly conservative life calculation results due to ignoring this process. The sum of the current damage depth and a preset correction value is used as the initial crack size for crack propagation analysis, and the initial crack propagation life is conservatively processed to obtain the target crack propagation life, reducing the error caused by directly equating minor damage to the initial crack calculation. The sum of the crack initiation life and the target crack propagation life yields the total low-cycle fatigue life of the damaged component, improving the reliability of low-cycle fatigue life assessment for aerospace components.

[0006] Optionally, a set of SN curves is established as follows: Multiple test pieces are fabricated using the same material as the aerospace part to be evaluated, and these test pieces are divided into two groups according to damage depth, with each group corresponding to a specific damage depth. Surface damage corresponding to the damage depth is pre-fabricated in the assessment area of ​​each test piece group. Each group of test pieces is further divided into at least two subgroups, with each subgroup corresponding to a specific load level. Low-cycle fatigue tests are conducted on each subgroup of test pieces at the corresponding load level, with multiple test pieces in each subgroup. Fatigue life data at the corresponding damage depth and load level are obtained based on the results of the low-cycle fatigue tests on each subgroup of test pieces. Damage SN curves corresponding to the two damage depths are then established based on the fatigue life data.

[0007] By adopting the above technical solution, test pieces are made of the same material as the aerospace parts to be evaluated. The test pieces are grouped according to the damage depth and the surface damage is pre-formed. Then, low-cycle fatigue tests are carried out in multiple subgroups according to the load level. Fatigue life data under the corresponding damage depth and load level are obtained. Then, damage SN curves corresponding to the two damage depths are established. This provides accurate basic data for subsequent interpolation calculation of crack initiation life based on the current damage depth and working stress, and improves the reliability of low-cycle fatigue life assessment of aerospace parts.

[0008] Optionally, fatigue life data at the corresponding damage depth and corresponding load level can be obtained based on the results of low-cycle fatigue tests of each subgroup of test specimens, including: calculating the test characteristic life corresponding to each subgroup of test specimens using a two-parameter Weibull distribution; and calculating the fatigue life at preset reliability and preset confidence levels based on the test characteristic life, specimen coefficient, confidence coefficient, and reliability coefficient, wherein the confidence coefficient and reliability coefficient are coefficients corresponding to the preset reliability and preset confidence levels, respectively.

[0009] By adopting the above technical solution, the test characteristic life corresponding to each subgroup of test specimens is calculated using a two-parameter Weibull distribution, which can accurately characterize the fatigue life distribution characteristics of the test specimens. Based on the test characteristic life, specimen coefficient, confidence coefficient, and reliability coefficient, the fatigue life under preset reliability and preset confidence is calculated, which can obtain fatigue life data that meets specific reliability and confidence requirements. This provides an accurate data basis for the subsequent establishment of damage SN curves, thereby improving the reliability of low-cycle fatigue life assessment of aerospace parts.

[0010] Optionally, a two-parameter Weibull distribution is used to calculate the test characteristic lifetime corresponding to each subgroup of test specimens, including: the target subgroup is any subgroup, the number of test specimens in the target subgroup is n, and the test characteristic lifetime corresponding to the target subgroup is calculated as follows: when all n test specimens in the target subgroup fail, When k out of n test pieces in the target subgroup fail and the remaining test pieces exceed their specified lifespan, ; where N i Let α be the test lifetime of the i-th test specimen in the target subgroup, α be the shape parameter of the two-parameter Weibull distribution corresponding to the sample material, and β be the test characteristic lifetime corresponding to the target subgroup.

[0011] By adopting the above technical solution, the test characteristic life of each subgroup of test pieces is accurately calculated using the two-parameter Weibull distribution and corresponding formulas. This provides a basis for subsequent calculation of fatigue life under preset reliability and preset confidence levels, making the fatigue life data obtained based on the test results more accurate. This, in turn, improves the accuracy of the damage SN curve and ultimately enhances the reliability of low-cycle fatigue life assessment of aerospace parts.

[0012] Optionally, based on the test characteristic life, sample coefficient, confidence coefficient, and reliability coefficient, calculate the fatigue life under preset reliability and preset confidence, including: calculating the fatigue life corresponding to the target subgroup in the following manner: N X / Y =β / (S T ×S C ×S R ), where N X / Y X represents the fatigue life corresponding to the target subgroup, Y represents the preset reliability, and S represents the preset confidence level. T S represents the sample coefficient. C S represents the confidence coefficient. R This represents the reliability coefficient.

[0013] By adopting the above technical solution, based on the test characteristic lifetime, sample coefficient, confidence coefficient, and reliability coefficient, according to formula N... X / Y =β / (S T ×S C ×S RCalculating the fatigue life corresponding to the target subgroup can accurately obtain the fatigue life under preset reliability and preset confidence levels, thereby providing a more accurate data basis for the subsequent establishment of damage SN curves. This helps to improve the reliability of low-cycle fatigue life assessment of aerospace parts damage and solves the problem of low reliability of assessment methods in related technologies.

[0014] Optionally, damage SN curves corresponding to two damage depths can be established based on fatigue life data, including: linearly fitting fatigue life data corresponding to at least two load levels at the same damage depth in a logarithmic coordinate system to obtain damage SN curves corresponding to the damage depth.

[0015] By adopting the above technical solution, the fatigue life data is linearly fitted in the logarithmic coordinate system to establish the damage SN curve, which can more accurately reflect the relationship between damage depth, load level and fatigue life, and provide a more accurate data basis for subsequent interpolation calculation of crack initiation life.

[0016] Optionally, the crack initiation life can be obtained by interpolation using the following formula: (Logd2-Logd) / (Logd2-Logd1)×(A1+B1×LogN0-A2-B2×LogN0)=σ-(A2+B2×LogN0); where d is the current damage depth, d1 is the first damage depth, d2 is the second damage depth, and d1<d<d2, σ is the working stress, A1 and B1 are the intercept and slope parameters of the SN curve corresponding to the first damage depth, respectively, A2 and B2 are the intercept and slope parameters of the SN curve corresponding to the second damage depth d2, respectively, and N0 is the crack initiation life.

[0017] By adopting the above technical solution, a set of SN curves are interpolated using a specific formula. Combined with the current damage depth and working stress of the aerospace parts to be evaluated, the crack initiation life can be accurately obtained, avoiding evaluation deviations caused by the "small crack effect" and ignoring the crack initiation process, and improving the reliability of low-cycle fatigue life assessment of aerospace parts.

[0018] Optionally, crack propagation analysis is performed based on the initial crack size to obtain the initial crack propagation life, including: calculating the stress intensity factor range at the crack tip using the following formula: ， Where, Δ K The stress intensity factor range is represented by F, the comprehensive correction factor is represented by Δσ, and the stress range is represented by Δσ. The crack length is represented by the following Paris formula, which is used to calculate the crack propagation rate based on the range of the stress intensity factor: Where C is the Paris coefficient reflecting the material's resistance to crack propagation, m is the Paris exponent reflecting the sensitivity of the crack propagation rate to the range of stress intensity factor, and N represents the number of cycles; the crack propagation life is calculated by integrating based on the crack propagation rate, according to the following three failure criteria, and the result corresponding to the failure criterion that is triggered first is taken as the initial crack propagation life: Fracture failure: Maximum stress intensity factor K at the crack tip max Exceeding the material's fracture toughness Kc, where, , σ max The maximum stress in the working stress spectrum of the part; static failure: net cross-sectional stress σ of the structure. n The material reaches the rheological stress, which is equal to the material's ultimate strength σ. ult With yield strength σ sy The average value; size failure: the crack length reaches the preset specified crack size.

[0019] By adopting the above technical solutions and calculating the stress intensity factor range at the crack tip using a specific formula, the stress change at the crack tip can be accurately quantified. The crack propagation rate can be calculated based on the stress intensity factor range using the Paris formula, which can scientifically reflect the speed of crack propagation. The crack propagation life can be calculated according to three failure criteria, and the result of the earliest triggering can be taken as the initial crack propagation life. This can comprehensively consider different failure conditions and obtain the initial crack propagation life more accurately, thereby making the low-cycle fatigue life assessment of the entire aerospace component more accurate and reliable.

[0020] Optionally, the above method further includes: for sample materials of the same material as the aerospace part to be evaluated, obtaining the fatigue life corresponding to the damage SN curve and the fatigue life corresponding to the no-damage SN curve under the same load level; dividing the fatigue life corresponding to the damage SN curve by the fatigue life corresponding to the no-damage SN curve to obtain the damage low-cycle fatigue reduction coefficient for the corresponding damage depth; and constructing the correspondence between damage depth and damage low-cycle fatigue reduction coefficient based on the damage low-cycle fatigue reduction coefficient corresponding to different damage depths.

[0021] By adopting the above technical solution, the fatigue life of damaged SN curves and undamaged SN curves under the same load level is obtained. The damage low-cycle fatigue reduction factor is calculated, and the correspondence between damage depth and damage low-cycle fatigue reduction factor is established. This can provide a basis for estimating the crack initiation life of aerospace parts made of the same type of material for which damage SN curves have not been established, and further improve the reliability and applicability of damage low-cycle fatigue life assessment of aerospace parts.

[0022] Optionally, the target aerospace part and the aerospace part to be evaluated are made of the same type of material. The above method also includes: when no damage SN curve corresponding to the target aerospace part is established, the crack initiation life of the target aerospace part is directly estimated based on the correspondence between damage depth and damage low-cycle fatigue reduction coefficient and the undamaged SN curve corresponding to the target aerospace part.

[0023] By adopting the above technical solution, when the damage SN curve corresponding to the target aerospace part has not been established, the crack initiation life of the target aerospace part can be directly estimated by using the correspondence between damage depth and damage low-cycle fatigue reduction factor and the undamaged SN curve corresponding to the target aerospace part. This avoids the complicated process of re-establishing the damage SN curve and improves the evaluation efficiency.

[0024] Optionally, the above method further includes: determining the permissible damage limit of the aerospace part to be evaluated based on the total low-cycle fatigue life of the aerospace part to be evaluated, the maintenance and inspection interval, and the preset safety margin.

[0025] By adopting the above technical solution, and combining the low-cycle fatigue life of the aircraft parts to be evaluated, the maintenance and inspection intervals, and the preset safety margins, the permissible damage limits of the aircraft parts can be scientifically and reasonably determined, thereby providing an important basis for the use and maintenance of aircraft parts and helping to ensure that aircraft parts serve within a safe range.

[0026] In a second aspect of this application, a low-cycle fatigue life assessment system for aerospace parts is also provided, for performing any of the aforementioned methods for assessing low-cycle fatigue life of aerospace parts, comprising: a first acquisition module for acquiring the current damage depth and working stress of the aerospace part to be assessed; a second acquisition module for acquiring a set of SN curves of a sample material of the same material as the aerospace part to be assessed, wherein the set of SN curves includes damage SN curves of the sample material at at least two different damage depths; a first obtaining module for obtaining crack initiation life by interpolating the set of SN curves based on the current damage depth and working stress; a processing module for taking the sum of the current damage depth and a preset correction value as the initial crack size, performing crack propagation analysis based on the initial crack size to obtain the initial crack propagation life, and performing conservative processing on the initial crack propagation life to obtain the target crack propagation life, wherein the conservative processing is used to represent the reduction processing of the initial crack propagation life; and a second obtaining module for summing the crack initiation life and the target crack propagation life to obtain the total low-cycle fatigue life of the aerospace part to be assessed.

[0027] In a third aspect of this application, an electronic device is also provided, including a memory and a processor, wherein a computer program is stored in the memory, and the processor executes the program to implement the method steps of any of the above claims.

[0028] In a fourth aspect of this application, a computer-readable storage medium is also provided, which stores instructions that, when executed, perform the method steps of any of the above claims.

[0029] In summary, one or more technical solutions provided in this application have at least the following technical effects or advantages: 1. The crack initiation process is considered to avoid overly conservative life calculation results caused by ignoring this process; the sum of the current damage depth and the preset correction value is used as the initial crack size for crack propagation analysis, and the initial crack propagation life is conservatively processed to obtain the target crack propagation life, reducing the error caused by directly equating minor damage to the initial crack calculation; the crack initiation life and the target crack propagation life are summed to obtain the total low-cycle fatigue life of the damage, which improves the reliability of low-cycle fatigue life assessment of aerospace parts. 2. Based on the test characteristic life, sample coefficient, confidence coefficient and reliability coefficient, the fatigue life under the preset reliability and preset confidence can be accurately obtained, thus providing a more accurate data basis for the subsequent establishment of damage SN curves and solving the problem of low reliability of evaluation methods in related technologies; 3. Establishing the correspondence between damage depth and damage low-cycle fatigue reduction factor can provide a basis for estimating the crack initiation life of aerospace parts made of the same type of material for which damage SN curves have not been established, and further improve the reliability and applicability of damage low-cycle fatigue life assessment of aerospace parts. Attached Figure Description

[0030] Figure 1 This is a flowchart of a method for assessing the low-cycle fatigue life of aircraft parts, provided in an embodiment of this application. Figure 2 This is a schematic diagram of the technical route for low-cycle fatigue life analysis of aerospace parts damage provided in the embodiments of this application; Figure 3 This is a schematic diagram of the test specimen provided in the embodiments of this application; Figure 4 This is a schematic diagram of the damage SN curve provided in the embodiments of this application; Figure 5 This is a schematic diagram of the reduction factor curve provided in the embodiments of this application; Figure 6 This is a schematic diagram of the initial crack size setting provided in the embodiments of this application; Figure 7 This is a schematic diagram of the crack propagation curve provided in the embodiments of this application; Figure 8 This is a schematic diagram illustrating the determination of permissible damage limits provided in an embodiment of this application; Figure 9This is a structural block diagram of a low-cycle fatigue life assessment system for aircraft parts damage provided in an embodiment of this application. Detailed Implementation

[0031] To enable those skilled in the art to better understand the technical solutions in this specification, the technical solutions in the embodiments of this specification will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of this application, and not all embodiments.

[0032] In the description of the embodiments of this application, the words "for example" or "for instance" are used to indicate examples, illustrations, or explanations. Any embodiment or design that is described as "for example" or "for instance" in the embodiments of this application should not be construed as being more preferred or advantageous than other embodiments or design options. Rather, the use of the words "for example" or "for instance" is intended to present the relevant concepts in a specific manner.

[0033] In the description of the embodiments of this application, the term "multiple" means two or more. Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the indicated technical features. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of that feature. The terms "comprising," "including," "having," and variations thereof all mean "including but not limited to," unless otherwise specifically emphasized.

[0034] This application provides a method for assessing the low-cycle fatigue life of aircraft components, referring to... Figure 1 , Figure 1 This is a flowchart of a method for assessing the low-cycle fatigue life of aerospace components, provided in an embodiment of this application, including the following steps: Step S101: Obtain the current damage depth and working stress of the aerospace part to be evaluated; Step S102: Obtain a set of SN curves of a sample material that is the same material as the aerospace part to be evaluated, which is established in advance. The set of SN curves includes the damage SN curves of the sample material at at least two different damage depths. Step S103: Based on the current damage depth and working stress, the crack initiation life is obtained by interpolating a set of SN curves. Step S104: The sum of the current damage depth and the preset correction value is used as the initial crack size. Crack propagation analysis is performed based on the initial crack size to obtain the initial crack propagation life. The initial crack propagation life is then conservatively processed to obtain the target crack propagation life. The conservative processing is used to indicate that the initial crack propagation life is reduced. Step S105: Sum the crack initiation life and the target crack propagation life to obtain the total low-cycle fatigue life of the aerospace part to be evaluated.

[0035] Through the above steps, the current damage depth and working stress of the aerospace component to be evaluated, as well as the damage SN curves of sample materials of the same material at at least two different damage depths, are obtained, providing basic data for subsequent life assessment. Interpolation calculation of a set of SN curves yields the crack initiation life, taking into account the crack initiation process and avoiding overly conservative life calculation results due to ignoring this process. The sum of the current damage depth and a preset correction value is used as the initial crack size for crack propagation analysis, and the initial crack propagation life is conservatively processed to obtain the target crack propagation life, reducing the error caused by directly equating minor damage to the initial crack calculation. The sum of the crack initiation life and the target crack propagation life yields the total low-cycle fatigue life of the damaged component, improving the reliability of the low-cycle fatigue life assessment of the aerospace component.

[0036] This embodiment provides a method for assessing the low-cycle fatigue life of aerospace components based on a two-stage model of "crack initiation + crack propagation". By constructing damage SN curves corresponding to at least two damage depths and interpolating them, the crack initiation life at a specific damage depth (such as the current damage depth) and working stress is accurately calculated. In the crack propagation stage, based on fracture mechanics theory, a preset correction value (such as 0.38 mm, or other values) is superimposed on the original damage as the initial crack size to avoid small crack effects. Crack propagation analysis is then performed and conservative reductions are applied. Finally, the low-cycle fatigue life of the two stages is added together to obtain the total low-cycle fatigue life. The current damage depth, measured on the surface of the component by non-destructive testing (such as eddy current testing, visual inspection with comparative test blocks), is the vertical depth of scratches, indentations, etc., measured in millimeters (mm). This is one of the input parameters of this assessment method. Working stress refers to the cyclic stress amplitude that the component experiences under extreme service conditions (such as maximum takeoff weight, maximum landing impact), measured in megapascals (MPa), which can usually be obtained through finite element analysis or measured strain data. The SN curve describes the relationship between stress level (S) and fatigue life (N). The horizontal axis represents the logarithm of life (LogN), and the vertical axis represents the logarithm of stress (Logσ), showing a linear relationship: Logσ = A + B × LogN. The damage SN curve is obtained through low-cycle fatigue testing on specimens with surface damage at a specific depth. Different damage depths correspond to different SN curves; the deeper the damage, the lower the curve position. A set of SN curves includes at least two SN curves corresponding to different damage depths (e.g., 0.1 mm and 0.3 mm, or other depths), used for subsequent interpolation calculations of the initiation life at any intermediate depth. These curves are obtained in advance through standard tests, not through on-site testing during the implementation of this method, demonstrating the practicality and efficiency of the method. The greater the damage depth, the shorter the initiation life under the same stress. Interpolation eliminates the need to test for every depth, significantly reducing costs. Conservative treatment refers to considering the dispersion of material properties and computational uncertainties, dividing the initial crack propagation life by a safety factor K (e.g., 2, or other coefficients greater than 1) to obtain the target crack propagation life. This is a standard engineering practice in the aerospace field to ensure safety. Total lifespan refers to the total number of cycles a component undergoes from its initial service life (either in good condition or with damage) to its eventual failure. This embodiment provides a comprehensive method for estimating the total lifespan of low-cycle fatigue fatigue, covering the entire fatigue failure process. It includes both the initiation stage, which consumes most of the lifespan, and the stable crack propagation stage, making it closer to actual physical behavior than the pure crack propagation method. In related technologies, the pure crack propagation method primarily ignores the crack initiation stage, leading to overly conservative lifespan assessments. Furthermore, the small crack effect results in significant discrepancies between theoretical calculations and physical behavior, resulting in insufficient reliability.This embodiment addresses the problems of unreliable small crack calculations, overly conservative results due to neglecting the initiation life, and inability to balance accuracy and safety caused by using only a single crack propagation method in related technologies. By using experimental curve interpolation, multi-coefficient correction, and two-stage coupled calculation, it achieves a reliable, accurate, and airworthiness compliant assessment of the low-cycle fatigue life of aerospace parts, effectively improving the accuracy of life assessment and its engineering practicality.

[0037] In an optional embodiment, a set of SN curves is established as follows: Multiple test pieces are fabricated using the same material as the aerospace part to be evaluated, and these test pieces are divided into two groups according to damage depth, with each group corresponding to a specific damage depth. Surface damage corresponding to the damage depth is pre-fabricated in the assessment area of ​​each group of test pieces. Each group of test pieces is further divided into at least two subgroups, with each subgroup corresponding to a specific load level. Low-cycle fatigue tests are performed on each subgroup of test pieces at the corresponding load level, wherein there are multiple test pieces in each subgroup. Fatigue life data at the corresponding damage depth and load level are obtained based on the results of the low-cycle fatigue tests on each subgroup of test pieces. Damage SN curves corresponding to the two damage depths are then established based on the fatigue life data.

[0038] In the above embodiments, test pieces are made of the same material as the aerospace parts to be evaluated, grouped according to damage depth and pre-fabricated with surface damage, and then divided into multiple subgroups according to load level for low-cycle fatigue testing. Fatigue life data under corresponding damage depth and load level are obtained, and damage SN curves corresponding to the two damage depths are established. This provides accurate basic data for subsequent interpolation calculation of crack initiation life based on the current damage depth and working stress, and improves the reliability of low-cycle fatigue life assessment of aerospace parts damage.

[0039] This embodiment provides a method for obtaining SN curves, proposing a low-cost, high-efficiency method for establishing damage SN curves through testing. The core of this method is a standardized process using the same material, grouped pre-existing damage, graded loads, and parallel testing of multiple specimens. Reliable life data is obtained through low-cycle fatigue tests with two damage depths, at least two load levels for each damage depth, and multiple specimens per load level, and then the damage SN curve is fitted. Specifically, the controlled variable method is used to obtain fatigue data covering typical working conditions through a small number of tests. First, two representative damage depths (e.g., 0.1mm and 0.3mm) are selected to cover the damage range commonly encountered in maintenance inspections. Second, for each damage depth, two different stress levels are selected (corresponding to the two ends of the low-cycle fatigue life range, such as 10...). 4 and 10 5Multiple specimens (e.g., 5 or more) are repeatedly tested at different damage depths. The test data are then statistically processed to obtain the characteristic life under each working condition. Finally, the two data points at each damage depth are connected in a logarithmic coordinate system to obtain the SN curve for that damage depth. This embodiment requires only two damage depths, two stress levels, and multiple specimens per subgroup to establish a complete damage SN curve system; interpolation can predict the initiation life at any intermediate depth and stress level, significantly reducing test costs and time. Pre-fabricated damage in the test area refers to pre-fabricating scratches or grooves of a set depth in the test section where stress is most concentrated on the specimen using machining methods; if specimen resources are limited, penetrating scratches (damage penetrating the width of the specimen) can be preferred, as their life results are more conservative and safer. Fatigue life data is the number of cycles N experienced by each specimen from the start of loading to failure (or exiting the test). i Taking a subgroup containing 5 test specimens as an example, for each subgroup at each damage depth and stress level (e.g., d1=0.1mm, σ=300MPa), the life data of the 5 specimens are recorded: N1, N2, N3, N4, and N5. The characteristic life can be calculated using a two-parameter Weibull distribution. The fatigue data is plotted in a double logarithmic coordinate system, with the horizontal axis being the logarithm of life N (LogN) and the vertical axis being the logarithm of stress σ (Logσ). For the same damage depth (e.g., d1=0.1mm), there are already two stress levels and their corresponding characteristic lives. In the logarithmic coordinate system, these two points determine a straight line with the equation Logσ=A+B×LogN. This embodiment ensures that the curve has material representativeness, statistical validity, and engineering applicability by controlling experimental variables, statistically analyzing multiple specimens, and applying graded loads. This provides a stable and accurate data foundation for the interpolation calculations in the aforementioned embodiment, improving the reliability and consistency of the overall machine life assessment.

[0040] In an optional embodiment, fatigue life data at the corresponding damage depth and corresponding load level are obtained based on the results of low-cycle fatigue tests of each subgroup of test specimens, including: calculating the test characteristic life corresponding to each subgroup of test specimens using a two-parameter Weibull distribution; and calculating the fatigue life at a preset reliability and a preset confidence level based on the test characteristic life, specimen coefficient, confidence coefficient, and reliability coefficient, wherein the confidence coefficient and reliability coefficient are coefficients corresponding to the preset reliability and preset confidence level, respectively.

[0041] In the above embodiments, the test characteristic life corresponding to each subgroup of test specimens is calculated using a two-parameter Weibull distribution, which can accurately characterize the fatigue life distribution characteristics of the test specimens. Based on the test characteristic life, specimen coefficient, confidence coefficient, and reliability coefficient, the fatigue life under preset reliability and preset confidence is calculated, which can obtain fatigue life data that meets specific reliability and confidence requirements. This provides an accurate data basis for the subsequent establishment of damage SN curves, thereby improving the reliability of low-cycle fatigue life assessment of aerospace parts.

[0042] This embodiment introduces a method for statistically processing fatigue test data based on the two-parameter Weibull distribution to derive fatigue life values ​​with engineering confidence from limited test data. The core principle is that fatigue life data inherently exhibits dispersion; even when testing the same batch of specimens at the same stress level, the measured lifespan will show a certain statistical distribution pattern. The two-parameter Weibull distribution is one of the most commonly used distribution models in fatigue life statistics, effectively describing the skewed distribution characteristics of fatigue life. The characteristic life β calculated using the Weibull distribution represents the lifespan value at which 63.2% of specimens fail, and is a characteristic parameter of this distribution. However, the characteristic life β corresponds to a 36.8% reliability (i.e., the average case). Aerospace engineering requires higher safety margins; therefore, it is necessary to further introduce reliability and confidence coefficients to correct the characteristic lifespan to a conservative lifespan value under specified reliability (e.g., 90%, 95%) and confidence (e.g., 90%, 95%). Directly using the arithmetic mean of test lives as the evaluation basis, without considering the statistical dispersion of fatigue data, may lead to risky evaluation results. By employing the Weibull distribution and reliability confidence level correction, a conservative lifespan value with clear statistical significance is obtained, meeting aviation safety standards. The Weibull distribution is a continuous probability distribution commonly used to describe lifespan data with skewed characteristics, such as fatigue life and material strength. Its probability density function is determined by two parameters: a shape parameter α (controlling the distribution morphology) and a scale parameter β (controlling the characteristic lifespan). The shape parameter α reflects the dispersion of lifespan data. The larger α is, the more concentrated the data; the smaller α is, the more dispersed the data. Different materials have different typical α values; for example, aluminum alloys: α≈4~6 (relatively concentrated data), titanium alloys: α≈3~5, high-strength steel: α≈2~4 (relatively dispersed data). In this embodiment, the characteristic lifespan of subgroup tests is first calculated from the results of multi-sample tests, and then corrected using the sample coefficient, confidence coefficient, and reliability coefficient to obtain the fatigue life at the specified reliability and confidence level that meets engineering safety requirements. This method addresses the issues of the susceptibility of directly using single-sample lifespan data to discreteness, the lack of statistical representativeness, and the inability to meet the confidence and reliability requirements for airworthiness. By fitting the fatigue life distribution law through Weibull distribution and using multi-coefficient joint correction, the method achieves the conversion from the sample test lifespan to the engineering usable lifespan, ensuring that the lifespan data has statistical validity, engineering conservatism, and airworthiness compliance. This provides standardized, reliable, and auditable input data for subsequent damage SN curve fitting.

[0043] In an optional embodiment, a two-parameter Weibull distribution is used to calculate the test characteristic lifetime corresponding to each subgroup of test specimens, including: the target subgroup is any subgroup, the number of test specimens in the target subgroup is n, and the test characteristic lifetime corresponding to the target subgroup is calculated as follows: when all n test specimens in the target subgroup fail... When k out of n test pieces in the target subgroup fail and the remaining test pieces exceed their specified lifespan, ; where N i Let α be the test lifetime of the i-th test specimen in the target subgroup, α be the shape parameter of the two-parameter Weibull distribution corresponding to the sample material, and β be the test characteristic lifetime corresponding to the target subgroup.

[0044] In the above embodiments, the test characteristic life of each subgroup of test pieces is accurately calculated using the two-parameter Weibull distribution and corresponding formulas, providing a basis for subsequent calculation of fatigue life under preset reliability and preset confidence levels. This makes the fatigue life data obtained based on the test results more accurate, thereby improving the accuracy of the damage SN curve and ultimately enhancing the reliability of low-cycle fatigue life assessment of aerospace parts.

[0045] This embodiment provides specific calculation rules for the characteristic life β under a two-parameter Weibull distribution. For the two common fatigue testing scenarios of complete failure and partial failure, unified and executable calculation formulas are provided. The characteristic life reflecting the overall performance of the test group is obtained by taking the weighted average of the test life to the power of α and then taking the square root of α. Two situations may occur during the test: all specimens fail during the test (complete failure); or some specimens reach the specified over-life without failure (partial failure). The calculation formula for β differs slightly in the two cases: for complete failure, data from all n specimens are used; for partial failure, data from only k failed specimens are used. By adopting the maximum likelihood estimation principle, taking into account both the inherent dispersion of the material (shape parameter α) and the actual test conditions, a stable, objective, and engineering-applicable characteristic life can be obtained regardless of whether all specimens fail or partially over-life. This provides an accurate statistical basis for subsequent reliability life calculations, making test data processing more standardized and more closely aligned with the real working conditions of aerospace fatigue testing. It solves the problem that in aerospace material testing, due to cost and cycle constraints, the sample size (n) is often very small, and the simple arithmetic mean cannot accurately reflect the fatigue life distribution characteristics. It also solves the data processing problem when some specimens do not break (exceed) because they have not reached the predetermined life.

[0046] In an optional embodiment, fatigue life under preset reliability and preset confidence is calculated based on test characteristic lifetime, sample coefficient, confidence coefficient, and reliability coefficient, including: calculating the fatigue life corresponding to the target subgroup in the following manner: N X / Y =β / (S T ×S C ×S R ), where N X / Y X represents the fatigue life corresponding to the target subgroup, Y represents the preset reliability, and S represents the preset confidence level. T S represents the sample coefficient. C S represents the confidence coefficient.R This represents the reliability coefficient.

[0047] In the above embodiments, based on the test characteristic lifetime, sample coefficient, confidence coefficient, and reliability coefficient, according to formula N X / Y =β / (S T ×S C ×S R Calculating the fatigue life corresponding to the target subgroup can accurately obtain the fatigue life under preset reliability and preset confidence levels, thereby providing a more accurate data basis for the subsequent establishment of damage SN curves. This helps to improve the reliability of low-cycle fatigue life assessment of aerospace parts damage and solves the problem of low reliability of assessment methods in related technologies.

[0048] This embodiment is based on the experimental characteristic lifetime β, and uses the sample coefficient S. T Confidence coefficient S C Reliability coefficient S R Triple correction converts the statistical lifetime of laboratory samples into fatigue lifetime N under specified reliability X and confidence level Y to meet aviation airworthiness requirements. X / Y Its core principle is that the test characteristic life only reflects the statistical results of the sample itself. It requires correction of the difference between the sample and the actual part through a sample coefficient, assurance of the statistical reliability of the evaluation results through a confidence coefficient, and assurance of structural safety through a reliability coefficient. This method solves the problems that test life alone cannot be directly used for engineering part evaluation, lacks a unified safety and statistical correction system, and is difficult to meet the high reliability and high confidence requirements of aviation. It achieves a standardized conversion from "test data" to "engineering usable life," providing a compliant, conservative, and auditable life point for the damage SN curve, supporting the accuracy and safety of overall aircraft life assessment.

[0049] Reliability X refers to the probability that a part will not fail within a given lifespan. For example, a reliability of 90% means that out of 100 parts, 90 will have a lifespan greater than this value, and 10 will have a lifespan less than this value. X is typically taken as 90%, 95%, or 99%, determined by aviation safety standards. Confidence level Y is based on the results of a finite sample test. We have Y% confidence that the true overall reliability is not lower than X. For example, a confidence level of 95% means that if the same test is repeated 100 times, the calculated result will fall within the confidence interval of the true value 95 times. Sample coefficient S T This takes into account the differences between the test specimen and the actual part (such as size effect, surface finish, stress gradient, etc.). A value of 1.3 is generally recommended; this is an empirical value. Confidence coefficient S C It is a coefficient related to the confidence level Y, the sample size n, and the shape parameter α. The higher the confidence level and the smaller the sample size, the higher the S value. C The larger the value. For example, when the confidence level is 95%, the sample size n=5, and α=3, S...C ≈1.218. Reliability coefficient S R This is a coefficient related to reliability X and shape parameter α. The higher the reliability, the higher the coefficient. R The larger the value. For example, when the reliability is 95% and α=3, S... R ≈2.7.

[0050] In an optional embodiment, damage SN curves corresponding to two damage depths are established based on fatigue life data, including: linearly fitting fatigue life data corresponding to at least two load levels at the same damage depth in a logarithmic coordinate system to obtain damage SN curves corresponding to the damage depth.

[0051] In the above embodiments, the damage SN curve is established by linearly fitting the fatigue life data in a logarithmic coordinate system, which can more accurately reflect the relationship between damage depth, load level and fatigue life, and provide a more accurate data basis for subsequent interpolation calculation of crack initiation life.

[0052] This embodiment obtains the damage-strain (SN) curve for the corresponding damage depth by linearly fitting conservative fatigue life data obtained at least two load levels at the same damage depth into a log-log coordinate system. Extensive fatigue test data shows that in a double log-log coordinate system (horizontal axis: logarithm of life LogN, vertical axis: logarithm of stress Logσ), the SN curve of metallic materials approximates a straight line, which can be described by the linear equation Logσ = A + B × LogN, where A is the intercept, and the slope B is negative, indicating that the greater the stress, the shorter the life. For the same damage depth, at least two life data points at different stress levels are needed to determine a straight line; two points are the minimum requirement, and more data points can be used to improve the fitting accuracy through linear regression. This invention addresses the problem of deriving a continuous, smooth, and statistically significant SN curve from a finite number of discrete test points (only 2-3 stress levels at each damage depth), providing a necessary functional model for subsequent lifetime interpolation under arbitrary stress. It also solves the problems of lacking standardized curve construction methods, high test costs due to excessive data points, and the inability to form a line from a single point, making it difficult to use for interpolation calculations. By fitting a straight line to at least two points, a stress-life one-to-one correspondence benchmark curve is formed, providing stable, accurate, and engineering-efficient basic data for interpolation calculations at arbitrary damage depths in the aforementioned embodiments. This makes the entire evaluation system feasible, low-cost, and highly reliable.

[0053] In an optional embodiment, the crack initiation life is obtained by interpolation according to the following formula: (Logd2-Logd) / (Logd2-Logd1)×(A1+B1×LogN0-A2-B2×LogN0)=σ-(A2+B2×LogN0); where d is the current damage depth, d1 is the first damage depth, d2 is the second damage depth, and d1<d<d2, σ is the working stress, A1 and B1 are the intercept parameter and slope parameter of the SN curve corresponding to the first damage depth, respectively, A2 and B2 are the intercept parameter and slope parameter of the SN curve corresponding to the second damage depth d2, respectively, and N0 is the crack initiation life.

[0054] In the above embodiments, a set of SN curves are interpolated using a specific formula. Combined with the current damage depth and working stress of the aerospace part to be evaluated, the crack initiation life can be accurately obtained, avoiding evaluation deviations caused by the "small crack effect" and ignoring the crack initiation process, and improving the reliability of low-cycle fatigue life assessment of aerospace parts.

[0055] This embodiment presents a crack initiation life calculation method based on bilinear interpolation. Its core principle is that, given the known SN curves (equations Logσ=A1+B1×LogN and Logσ=A2+B2×LogN) corresponding to two damage depths d1 and d2, for any damage depth d between d1 and d2, its SN curve in a bilinear coordinate system can be obtained by linearly interpolating the parameters of the two known curves. The above formula is the mathematical expression of this interpolation idea; it establishes a functional relationship between the logarithm of the damage depth, the working stress σ, and the crack initiation life N0 to be determined. N0 can be obtained by solving this equation. Here, N... O This method also targets the lifetime corresponding to reliability X and confidence level Y. It addresses the problems in practical engineering where damage depths vary widely, making it impossible to establish individual curves for each depth and resulting in insufficient accuracy due to estimations alone. It eliminates the need for finite element modeling and additional experiments, directly calculating accurate, conservative, and repeatable crack initiation lifetimes through formulas. This enables rapid and high-precision assessment of damage across the entire size range, perfectly matching the aforementioned two-stage lifetime model and improving engineering practicality and assessment efficiency.

[0056] In an optional embodiment, crack propagation analysis is performed based on the initial crack size to obtain the initial crack propagation life, including: calculating the stress intensity factor range at the crack tip according to the following formula: ， Where, Δ K The stress intensity factor range is represented by F, the comprehensive correction factor is represented by Δσ, and the stress range is represented by Δσ. The crack length is represented by the following Paris formula, which is used to calculate the crack propagation rate based on the range of the stress intensity factor: Where C is the Paris coefficient reflecting the material's resistance to crack propagation, m is the Paris exponent reflecting the sensitivity of the crack propagation rate to the range of stress intensity factor, and N represents the number of cycles; the crack propagation life is calculated by integrating based on the crack propagation rate, according to the following three failure criteria, and the result corresponding to the failure criterion that is triggered first is taken as the initial crack propagation life: Fracture failure: Maximum stress intensity factor K at the crack tip max Exceeding the material's fracture toughness Kc, where, , σ max The maximum stress in the working stress spectrum of the part; static failure: net cross-sectional stress σ of the structure. n The material reaches the rheological stress, which is equal to the material's ultimate strength σ. ult With yield strength σ sy The average value; size failure: the crack length reaches the preset specified crack size.

[0057] In the above embodiments, the stress intensity factor range at the crack tip is calculated according to a specific formula, which can accurately quantify the stress change at the crack tip; the crack propagation rate is calculated based on the stress intensity factor range using the Paris formula, which can scientifically reflect the speed of crack propagation; the crack propagation life is calculated according to three failure criteria respectively, and the result of the earliest triggering is taken as the initial crack propagation life, which can comprehensively consider different failure conditions and obtain the initial crack propagation life more accurately, thereby making the low-cycle fatigue life assessment of the entire aerospace component damage more accurate and reliable.

[0058] This embodiment starts with the initial crack size, first calculating the range of the stress intensity factor at the crack tip, then obtaining the crack propagation rate using the Paris formula, and finally obtaining the crack propagation life through integration. Simultaneously, it judges three criteria in parallel: fracture failure, static failure, and dimensional failure, taking the life corresponding to the earliest failure as the initial crack propagation life. The principle is that the crack propagation behavior under cyclic loading follows the Paris formula, and the crack propagation rate d... l / dN is proportional to the m-th power of the stress intensity factor range ΔK. ΔK is determined by the crack length. l The stress range Δσ and the geometric correction factor F are determined, and the expression is: From the initial crack length l 0 integral to critical crack length l c The crack propagation life can then be obtained. Critical crack length l c is determined by the minimum value among the three failure criteria: fracture failure (K) maxThe failure of a component is determined by three conditions: reaching the material fracture toughness Kc, static failure (net cross-sectional stress reaching the flow stress), and dimensional failure (reaching the preset geometric boundary). The first of these conditions to be met is considered a component failure. This approach addresses the problems of traditional single-crack propagation methods, which fail to consider multiple failure modes, rely solely on fracture criteria (prone to bias), neglect structural load-bearing capacity and airworthiness size limitations, and fail to balance safety and engineering realities. By using multiple constraints and selecting the shortest conservative lifespan, the assessment results more closely reflect real structural failure behavior, while avoiding calculation biases caused by small crack effects. This provides a reliable, safe, and airworthiness-compliant crack propagation lifespan for total lifespan calculations. This embodiment comprehensively considers three failure modes: fracture, static strength, and geometric boundary, using the most stringent one as the control criterion to ensure the safety of the assessment results. The comprehensive correction factor F, material parameters C and m in the above formulas can all be obtained from handbooks.

[0059] Example: Suppose the fuselage and wing connection joint of a certain type of transport aircraft to be evaluated is made of 7075-T6 aluminum alloy. Inspection revealed a scratch on the surface of the joint with a measured depth of d=0.22mm. Through finite element analysis, the cyclic stress amplitude of the joint under maximum takeoff conditions is σ=300MPa. The SN curves for damage depths d1=0.1mm and d2=0.3mm have been pre-established for this material.

[0060] The complete process of adopting the scheme of this embodiment includes: (1) obtaining input parameters: current damage depth d=0.22mm, working stress σ=300MPa; (2) calling the pre-established SN curve: the SN curve equations of two damage depths are known (obtained through previous experiments), d1=0.1mm: Logσ=3.20-0.12×LogN (A1=3.20, B1=-0.12), d2=0.3mm: Logσ=3.05-0.13×LogN (A2=3.05, B2=-0.13); (3) interpolating to calculate crack initiation life: substituting into the above formula to solve N0, N0≈41700 cycles; (4) crack propagation analysis: initial crack size l 0 = d + 0.38 = 0.22 + 0.38 = 0.60 mm, material parameter (7075-T6) C = 2.0 × 10 -11 m=3.2, F=1.12, Δσ=300MPa; fracture toughness The critical crack size at fracture was calculated. l c ≈3.96mm; from l 0 = 0.60mm integrated to l c =3.96mm: ≈9,200 iterations; Conservative approach: take a safety factor K=2, N CG / K=9,200 / 2=4,600 cycles; (5) Calculate the total lifetime: N total =41,700 + 4,600 = 46,300 cycles.

[0061] In an optional embodiment, the method further includes: for a sample material of the same material as the aerospace part to be evaluated, obtaining the fatigue life corresponding to the damage SN curve and the fatigue life corresponding to the no-damage SN curve under the same load level; dividing the fatigue life corresponding to the damage SN curve by the fatigue life corresponding to the no-damage SN curve to obtain the damage low-cycle fatigue reduction coefficient corresponding to the damage depth; and constructing a correspondence between damage depth and damage low-cycle fatigue reduction coefficient based on the damage low-cycle fatigue reduction coefficients corresponding to different damage depths.

[0062] In the above embodiments, the fatigue life of damaged SN curves and undamaged SN curves under the same load level is obtained, and the damage low-cycle fatigue reduction factor is calculated. The correspondence between damage depth and damage low-cycle fatigue reduction factor is established, which can provide a basis for estimating the crack initiation life of aerospace parts of the same type of material for which damage SN curves have not been established, and further improve the reliability and applicability of damage low-cycle fatigue life assessment of aerospace parts.

[0063] This embodiment establishes a correlation between damage depth and life reduction by comparing the fatigue lives of damaged and undamaged materials under the same stress level. The core principle is to quantify the impact of damage on life as a dimensionless reduction coefficient. The principle is that under the same load, the ratio of damaged life to undamaged life can stably reflect the degree of damage reduction. This coefficient has material universality and can be used across materials of the same type. This method solves the problems of lacking rapid methods for estimating damage life, inability to assess new materials / parts with undamaged SN curves, high experimental costs, and long cycles. By establishing a correlation of reduction coefficients, it achieves rapid life prediction under undamaged SN curve conditions, expanding the applicability of the assessment method, improving engineering application efficiency, and providing a unified and concise quantitative tool for damage assessment across materials and parts. For example, for the same base material (such as aluminum alloy), different grades (such as 7075, 2024, and 6061) may have different specific performance parameters, but the reduction law of damage on fatigue life is similar. Suppose that the Z-values ​​of a certain material (e.g., 7075) at several damage depths are obtained through experiments. A curve showing the relationship between Z and damage depth d can be plotted, and this curve can be extended to other grades in the same series (e.g., 2024, 6061). The reduction factor Z = damage life ÷ no-damage life. Z is a dimensionless coefficient less than 1, reflecting how much the life is reduced by damage. Z is calculated for multiple damage depths (d1, d2…), forming a d-Z curve / table, which can be used for interpolation to obtain the reduction factor corresponding to any damage depth. For example, for TC4 titanium alloy, assuming the stress level σ = 450 MPa, the no-damage life (smooth specimen) is found to be: N_ND = 28,000 cycles, and the damage depth d = 0.2 mm. The calculated damage life is: N X / Y =17,600 times, reduction factor Z = 17600 ÷ 28000 = 0.629; then take d = 0.1mm and d = 0.3mm and calculate similarly, we get: d = 0.1mm → Z1 ≈ 0.78, d = 0.2mm → Z2 ≈ 0.63, d = 0.3mm → Z3 ≈ 0.51, finally establishing the correspondence between damage depth d and reduction factor Z. Through this embodiment, only a complete damage SN curve test is needed for a typical material to establish the reduction factor curve; for other grades of materials in the same series, only their undamaged SN curve is known, and multiplying it by the reduction factor is needed to calculate the damage SN curve, greatly reducing the amount of testing.

[0064] In an optional embodiment, the target aerospace part and the aerospace part to be evaluated are made of the same type of material. The method further includes: when no damage SN curve corresponding to the target aerospace part is established, the crack initiation life of the target aerospace part is directly estimated based on the correspondence between damage depth and damage low-cycle fatigue reduction coefficient and the undamaged SN curve corresponding to the target aerospace part.

[0065] In the above embodiments, when the damage SN curve corresponding to the target aerospace part is not established, the crack initiation life of the target aerospace part can be directly estimated by using the correspondence between damage depth and damage low-cycle fatigue reduction coefficient and the undamaged SN curve corresponding to the target aerospace part. This avoids the complicated process of re-establishing the damage SN curve and improves the evaluation efficiency.

[0066] This embodiment is based on the core principle that damage reduction laws can be transferred among similar materials. When a damage SN curve has not been established for the target aerospace part, the damage depth-reduction coefficient correspondence constructed in the previous embodiment is used, combined with the target part's own undamaged SN curve, to directly and quickly estimate its crack initiation life. The principle is that similar metallic materials (such as titanium alloys and high-strength structural steel) have similar damage sensitivity characteristics, and the reduction coefficient has cross-grade universality. Existing reduction relationships for reference materials can be applied to new materials and new structural components without the need for extensive damage fatigue testing. This method solves the problems of being unable to evaluate new materials / structures without a damage SN curve, high testing costs, long testing cycles, and lack of basis for emergency field assessments. It achieves rapid, low-cost, and reliable damage life estimation, significantly expanding the applicability of the entire assessment method while taking into account both engineering practicality and aerospace airworthiness safety requirements. In practical applications, for target materials lacking damage test data (such as the aforementioned target aerospace parts), it is sufficient to obtain their undamaged SN curve (which can be obtained from material handbooks or determined through a small number of experiments), and then multiply it by the reduction factor Z(d) established for typical materials. This yields the damage fatigue life of the target material under the same damage depth and working stress. This method is essentially a similarity extrapolation technique, based on the engineering experience that "the relative impact of damage on fatigue performance is not significantly related to the specific material grade." Through this embodiment, it is unnecessary to conduct complete damage fatigue tests on every material. By utilizing existing reduction factor curves and the undamaged SN curve of the target material, damage life assessment results can be obtained quickly and economically, greatly expanding the applicability and engineering practicality of the method.

[0067] In an optional embodiment, the method further includes: determining the permissible damage limit of the aerospace part to be evaluated based on the total low-cycle fatigue life of the aerospace part to be evaluated, the maintenance and inspection interval, and the preset safety margin.

[0068] In the above embodiments, by combining the total low-cycle fatigue life of the aircraft part to be evaluated, the maintenance and inspection interval, and the preset safety margin, the permissible damage limit of the aircraft part can be scientifically and reasonably determined, thereby providing an important basis for the use and maintenance of the aircraft part and helping to ensure that the aircraft part is in service within a safe range.

[0069] This embodiment, based on the principles of continuous airworthiness and safety margin control for aerospace structures, combines the calculated total low-cycle fatigue life of the damaged component, maintenance and inspection intervals, and preset safety margins to reverse-determine the maximum acceptable damage size of the component, i.e., the Allowable Damage Limit (ADL). Its core logic is that allowable damage must ensure that, within one inspection cycle, even if there are missed inspections, the remaining fatigue life of the component is still greater than the inspection interval multiplied by the safety margin, ensuring that fatigue failure will not occur before the next inspection. This method solves the problems of traditional allowable damage limits relying heavily on experience, lacking forward calculation methods, being unable to quantitatively determine for low-cycle fatigue scenarios, and struggling to balance safety and economy. Through quantitative constraints of life, inspection cycle, and safety margin, it provides scientific, verifiable, and airworthiness compliant allowable damage limits, filling the gap in forward design of allowable damage for aerospace low-cycle fatigue structures, and supporting the development of maintenance manuals and field damage rejection standards.

[0070] Specifically, a correlation is established between the damage depth of the aerospace part to be evaluated and the total low-cycle fatigue life of the damage; based on the maintenance and inspection interval and the preset safety margin, the minimum required life is determined; based on the correlation between the damage depth and the total low-cycle fatigue life of the damage, the maximum damage depth that meets the minimum required life is determined, and the maximum damage depth is determined as the allowable damage limit of the aerospace part to be evaluated.

[0071] For example, the part to be evaluated is a titanium alloy aircraft connector, with an inspection interval of 15,000 takeoffs and landings and a safety margin (considering missed inspections) of 2 times. The requirement is: total life ≥ 2 × 15,000 = 30,000 takeoffs and landings. This method can be used to calculate: damage depth 0.15mm → total life 34,000 cycles, damage depth 0.175mm → total life 30,000 cycles, damage depth 0.2mm → total life 24,000 cycles. The comparison requirement is ≥ 30,000 cycles. The maximum depth meeting the condition is 0.175mm; therefore, the allowable damage limit (ADL) for this part is 0.175mm. For this part, with a scratch depth < 0.175mm at a 15,000-cycle inspection interval, repair is not required and it can be released.

[0072] The following description, in conjunction with specific embodiments, illustrates a method for analyzing the low-cycle fatigue life of aircraft parts damaged by damage. This method combines experimental and analytical approaches, and the technical route is as follows: Figure 2 As shown. This technology is applied to the maintenance and inspection of aerospace components and ongoing airworthiness work, including but not limited to the establishment of Allowable Damage Limits (ADL) values ​​in manuals and the assessment of remaining service life due to field damage. Applicable components include metal structural parts of aircraft fuselages, landing gear structures, and engines. Applicable damage types include scratches, abrasions, nicks, and similar damage.

[0073] The low-cycle fatigue process of a component can be divided into two stages: the crack initiation stage and the crack propagation stage.

[0074] During the crack initiation stage, the damage SN curve of the material is obtained experimentally, and the crack propagation life N is then calculated based on the working stress of the part. The test specimen is a square cross-section plate specimen, as shown in the schematic diagram. Figure 3 As shown. The steps are as follows: ① Select the type of damage to be analyzed, including scratches, notches, etc. If test specimen resources are limited, prioritize penetrating scratch damage, as the life results will be more conservative.

[0075] ② In the test section of the test piece, pre-fabricate damage, which is divided into two different damage depths to cover the common damage depth range of maintenance and inspection. The machining method is selected first for pre-fabrication.

[0076] ③ Using stress control methods, low-cycle fatigue tests were conducted. Two sets of tests with different load levels were performed at each damage depth (i.e., a total of four sets). The lifespan of the test specimens was [10...]. 4 10 5 ] times the loop and [1.5×10 5 5×10 5 The optimal range is within the cycle range, with a minimum of 5 valid test specimens per set for each damage depth and load level.

[0077] ④ Calculate the test characteristic life β. When all n test pieces fail, β is calculated according to formula (1): ……(1); When k out of all n test specimens fail (the remaining test specimens reach their specified lifespan, e.g., 5 × 10⁻⁶), the failure occurs. 5 (In the next cycle), β is calculated according to formula (2): ……(2); In equations (1) and (2), α is the shape parameter of the two-parameter Weibull distribution, which takes different values ​​depending on the material type, such as 3 for titanium alloy.

[0078] ⑤ Considering the dispersion of material fatigue properties, calculate the lifetime NX / Y under a certain reliability X and confidence level Y, as shown in formula (3).

[0079] ...(3); In the formula, S T For the sample coefficient, a value of 1.3 is recommended; S C S is the confidence coefficient, which varies with confidence level Y, number of samples n, and shape parameter α. For example, when the confidence level is 95%, the number of samples is 5, and the shape parameter α is 3, S C Take 1.218; S RS is the reliability coefficient, which varies with reliability X and shape parameter α. For example, when the reliability is 95% and the shape parameter α is 3, S... R Take 2.7.

[0080] ⑥ The two points N are calculated from the test results of two sets of different loads at a certain damage depth. X / Y Connecting the lines yields the SN curve at this damage depth, with the horizontal axis representing the logarithm of life (LogN) and the vertical axis representing the logarithm of stress (Logσ), as shown below. Figure 4 As shown. The SN curve for a larger damage depth should be below the smaller one, and the equation of the straight line is as shown in formula (4).

[0081] ……(4); In the formula, A is the intercept parameter and B is the slope parameter.

[0082] ⑦ For surface damage with a depth of d found during inspection, the NX / Y of the two damage SN curves with damage depths of d1 and d2 (d2>d1) is calculated by interpolating the logd parameter. This is the conservative crack initiation life, as shown in formula (5).

[0083] ……(5); In the formula, σ represents the working stress, and N0 is the lifetime corresponding to reliability X and confidence Y. This formula is applicable to

[10] . 4 10 5 Calculation of lifetime number of cycles.

[0084] In addition, by comparing the experimental data with the material SN curve (without damage), the damage low-cycle fatigue reduction factor Z under a certain damage depth d can be obtained, as shown in formula (6).

[0085] ...(6); In the formula, N ND This is the no-damage fatigue life with confidence level Y and reliability X obtained from the SN curve (no damage) of the same material. A schematic diagram of the reduction factor curve is obtained based on the reduction factor at different depths d, as shown below. Figure 5 As shown.

[0086] When damage SN curves for a specific material grade are unavailable, the damage crack initiation life under a certain damage depth d and working stress can be estimated using the reduction factor curves accumulated from previous tests and the material SN curves (without damage). For example, the damage fatigue life of parts using Ti17 material can be estimated using the reduction factor curves obtained from tests on TC4 and Ti6242 materials.

[0087] ⑧ After obtaining the crack initiation life, the crack propagation life can be calculated. It can be assumed that after crack initiation, a crack with a size equal to the original damage size (damage depth, damage half-length, damage radius) + 0.38 mm is formed. Figure 6 As shown, this is used as the initial crack size, and the crack propagation life is obtained through crack propagation analysis.

[0088] During crack propagation analysis, the range of stress intensity factor ΔK is calculated using formula (7).

[0089] ……(7); In the formula, F is a comprehensive correction factor, which is related to geometric parameters and loading method; Δσ is the stress range; l The crack length is given.

[0090] The crack propagation rate is calculated using the Paris formula, as shown in formula (8).

[0091] ……(8); In the formula, C and m are material parameters. A schematic diagram of the crack propagation curve is shown below. Figure 7 As shown. There are three criteria for crack propagation failure; failure is considered to have occurred if any one of these criteria is met.

[0092] One type is fracture failure, specifically the stress intensity factor K at the crack tip. max Exceeding the material's fracture toughness Kc, K max The calculation formula is shown in formula (9), where Kc is the material parameter.

[0093] ...(9); In the formula, This represents the maximum stress in the working stress spectrum of the part.

[0094] The second is static failure, which means that the net cross-sectional stress of the analyzed structure reaches the flow stress (the median of the material's fracture strength and yield strength), as shown in formula (10); ……(10); In the formula, The net cross-sectional stress can be calculated from the working stress and the remaining load-bearing area of ​​the part. The ultimate strength of the material, The yield strength of the material.

[0095] Third, to achieve the specified crack size, such as when the skin crack reaches two spans in length.

[0096] Crack propagation life N CG Material dispersion also needs to be considered, and the coefficient K (generally taken as 2) should be divided to obtain the conservative crack propagation life N.CG / K (corresponding to the aforementioned target crack propagation life), as shown in formula (11).

[0097] ……(11); Ultimately, the low-cycle fatigue life of a component is equal to the conservatively estimated crack initiation life N. X / Y +Conservative crack propagation life N CG / K.

[0098] This method for analyzing the low-cycle fatigue life of aircraft parts damage has applications including, but not limited to, assessing the remaining life of damage in the field and establishing permissible damage limits in continuing airworthiness documentation manuals. The former's principle is that for damage of depth *d* found in the field, the remaining low-cycle fatigue life is obtained through the above analysis, assessing how long it can be safely used without repair. The latter's principle is to analyze the relationship between damage depth *d* and low-cycle fatigue life, and based on the part's inspection interval, considering the possibility of missed inspections, determine what size of damage is permissible for repair-free use under a certain maintenance procedure. For example, if a part's inspection interval is 15,000 flight cycles, considering the possibility of missed inspections, the remaining low-cycle fatigue life is required to be greater than twice the inspection interval, i.e., 30,000 flight cycles. Based on the scratch damage depth-low-cycle fatigue life relationship curve for this part, the permissible scratch damage depth is determined to be 0.175 mm. See the schematic diagram below. Figure 8 As shown.

[0099] Compared with related technologies, the embodiments of this application have at least the following technical effects: Compared to directly performing crack propagation analysis, this method considers the crack initiation life of the damage, which greatly increases the remaining damage life and solves the problems of the former being too conservative and the low accuracy of small crack calculation. The analysis method that combines experimentation and calculation yields a higher level of reliability for remaining lifetime compared to purely calculation-based methods. Material damage SN curve testing requires fewer test specimens and has a lower cost; it can predict the residual crack initiation life under different damage depths and working stresses. A reduction factor curve for the initiation life of low-cycle fatigue cracks was formed, which can be extended to various grades of aluminum alloys, titanium alloys and steel materials used in the aerospace field. The calculation method is simple and efficient. It eliminates the need to use the finite element method to calculate the crack stress intensity factor, saving the complex steps of modeling and iterative calculations, and takes into account the material dispersion and performs conservative treatment. In terms of applications, for aerospace parts that are mainly affected by low-cycle fatigue, this method fills the gap in the forward calculation of their allowable damage limits.

[0100] In summary, the method of this embodiment has the advantages of low experimental difficulty and cost, simple calculation steps, applicability to a variety of materials, appropriate conservatism, and good engineering practicality.

[0101] This application also provides a low-cycle fatigue life assessment system for aircraft parts, used to perform the low-cycle fatigue life assessment method for aircraft parts damage in any of the foregoing embodiments, such as... Figure 9 As shown, the system includes: The first acquisition module is used to acquire the current damage depth and working stress of the aerospace part to be evaluated; The second acquisition module is used to acquire a set of SN curves of a sample material that is the same material as the aerospace part to be evaluated, which is pre-established. The set of SN curves includes the damage SN curves of the sample material at at least two different damage depths. The first acquisition module is used to obtain the crack initiation life by interpolating a set of SN curves based on the current damage depth and working stress. The processing module is used to take the sum of the current damage depth and the preset correction value as the initial crack size, perform crack propagation analysis based on the initial crack size to obtain the initial crack propagation life, and perform conservative processing on the initial crack propagation life to obtain the target crack propagation life. The conservative processing is used to represent the reduction processing of the initial crack propagation life. The second acquisition module is used to sum the crack initiation life and the target crack propagation life to obtain the total low-cycle fatigue life of the aerospace part to be evaluated.

[0102] It should be noted that the devices or systems provided in the above embodiments are only illustrated by the division of the above functional modules. In practical applications, the above functions can be assigned to different functional modules as needed, that is, the internal structure of the device can be divided into different functional modules to complete all or part of the functions described above. In addition, the system and method embodiments provided in the above embodiments belong to the same concept. Other device or system embodiments correspond to the aforementioned method embodiments. Other technical features are described in the previous embodiments and will not be repeated here.

[0103] This application also provides a computer-readable storage medium storing instructions that, when executed, perform the steps of any of the methods described above.

[0104] In one exemplary embodiment, the aforementioned computer-readable storage medium may include, but is not limited to, various media capable of storing computer programs, such as a USB flash drive, read-only memory (ROM), random access memory (RAM), portable hard disk, magnetic disk, or optical disk.

[0105] In the above embodiments, the descriptions of each embodiment have different focuses. For parts not described in detail in a certain embodiment, please refer to the relevant descriptions in other embodiments.

[0106] In the various embodiments provided in this application, it should be understood that the disclosed apparatus or system can be implemented in other ways. For example, the apparatus or system embodiments described above are merely illustrative; for instance, the division of units is only a logical functional division, and in actual implementation, there may be other division methods. For example, multiple units or components may be combined or integrated into another system, or some features may be ignored or not executed. Furthermore, the coupling or direct coupling or communication connection shown or discussed may be through some service interface; the indirect coupling or communication connection between apparatuses or units may be electrical or other forms.

[0107] The above description is merely an exemplary embodiment of this disclosure and should not be construed as limiting the scope of this disclosure. Any equivalent changes and modifications made in accordance with the teachings of this disclosure shall still fall within the scope of this disclosure. Other embodiments of this disclosure will be readily apparent to those skilled in the art upon consideration of the disclosure herein.

[0108] This application is intended to cover any variations, uses, or adaptations of this disclosure that follow the general principles of this disclosure and include common knowledge or customary techniques in the art that are not described in this disclosure.

Claims

1. A method for assessing low cycle fatigue life of damage in an aerospace part, characterized by, The method includes: Obtain the current damage depth and operating stress of the aerospace component to be evaluated; Obtain a set of SN curves for a sample material of the same material as the aerospace part to be evaluated, wherein the set of SN curves includes the damage SN curves of the sample material at at least two different damage depths; Based on the current damage depth and the working stress, the crack initiation life is obtained by interpolating the set of SN curves. The sum of the current damage depth and the preset correction value is used as the initial crack size. Crack propagation analysis is performed based on the initial crack size to obtain the initial crack propagation life. The initial crack propagation life is then conservatively processed to obtain the target crack propagation life. The conservative processing is used to indicate that the initial crack propagation life is reduced. The total low-cycle fatigue life of the aerospace component to be evaluated is obtained by summing the crack initiation life and the target crack propagation life.

2. The method of claim 1, wherein, The set of SN curves was established in the following way: Multiple test pieces are made from the same material as the aerospace part to be evaluated. The test pieces are divided into two groups according to the damage depth. Each group of test pieces corresponds to a damage depth. Surface damage corresponding to the damage depth is pre-made in the assessment area of ​​each group of test pieces. Each group of test specimens is divided into at least two subgroups, and each subgroup of test specimens corresponds to a load level. Low-cycle fatigue tests are performed on each subgroup of test specimens using the corresponding load level. The number of test specimens in each subgroup is multiple. Based on the results of low-cycle fatigue tests on the test specimens of each subgroup, fatigue life data at the corresponding damage depth and corresponding load level were obtained respectively. Damage SN curves corresponding to two damage depths were established based on fatigue life data.

3. The method of claim 2, wherein, Based on the results of low-cycle fatigue tests on the test specimens of each subgroup, fatigue life data at the corresponding damage depth and corresponding load level were obtained, including: The test characteristic life of each subgroup of test specimens was calculated using a two-parameter Weibull distribution. Based on the test characteristic life, sample coefficient, confidence coefficient, and reliability coefficient, the fatigue life under the preset reliability and preset confidence is calculated, wherein the confidence coefficient and the reliability coefficient are coefficients corresponding to the preset reliability and the preset confidence, respectively.

4. The method of claim 3, wherein, The characteristic lifetime of each subgroup of test specimens was calculated using a two-parameter Weibull distribution, including: The target subgroup is any subgroup, and the number of test specimens in the target subgroup is n. The test characteristic lifetime corresponding to the target subgroup is calculated as follows: When all n test pieces in the target subgroup are completely destroyed, ; when k of n test pieces in the target subset fail and the rest of the test pieces exceed a specified lifetime, ; Where, N i Let α be the test lifetime of the i-th test specimen in the target subgroup, α be the shape parameter of the two-parameter Weibull distribution corresponding to the sample material, and β be the test characteristic lifetime corresponding to the target subgroup.

5. The method of claim 4, wherein, Based on the test characteristic life, sample coefficient, confidence coefficient, and reliability coefficient, calculate the fatigue life under preset reliability and preset confidence, including: The fatigue life corresponding to the target subgroup is calculated as follows: N X / Y = β / (S T × S C × S R ), Where, N X / Y X represents the fatigue life corresponding to the target subgroup, Y represents the preset reliability, and S represents the preset confidence level. T S represents the sample coefficient. C S represents the confidence coefficient. R This represents the reliability coefficient.

6. The method of claim 1, wherein, The crack initiation life can be obtained by interpolation using the following formula: (Logd2-Logd) / (Logd2-Logd1)×(A1+B1×LogN0-A2-B2×LogN0)=σ-(A2+B2×LogN0); Wherein, d is the current damage depth, d1 is the first damage depth, d2 is the second damage depth, and d1 < d < d2, σ is the working stress, A1 and B1 are the intercept parameter and slope parameter of the SN curve corresponding to the first damage depth, respectively, A2 and B2 are the intercept parameter and slope parameter of the SN curve corresponding to the second damage depth d2, respectively, and N0 is the crack initiation life.

7. The method of claim 1, wherein, Crack propagation analysis is performed based on the initial crack size to obtain the initial crack propagation life, including: The stress intensity factor range at the crack tip is calculated using the following formula: ， Where, Δ K The stress intensity factor range is represented by F, the comprehensive correction factor is represented by Δσ, and the stress range is represented by Δσ. Indicates the crack length; The crack propagation rate is calculated using the following Paris formula based on the stress intensity factor range: where C is a Paris coefficient reflecting the crack propagation resistance of the material, m is a Paris exponent reflecting the sensitivity of the crack propagation rate to the stress intensity factor range, and N denotes the number of cycles. Based on the crack propagation rate, an integral calculation is performed, and the corresponding crack propagation lifetime is calculated according to the following three failure criteria, with the result corresponding to the first triggered failure criterion being taken as the initial crack propagation lifetime: Fracture failure: maximum stress intensity factor K at crack tip max exceeds the material fracture toughness Kc, where, , σ max is the maximum stress in the part's working stress spectrum; Static failure: net cross-sectional stress σ of the structure n The material reaches the rheological stress, where the rheological stress is the material's ultimate strength σ. ult With yield strength σ sy The average value; Size failure: The crack length reaches the preset specified crack size.

8. The method of claim 1, wherein, The method further includes: For sample materials made of the same material as the aerospace part to be evaluated, the fatigue life corresponding to the damage SN curve and the fatigue life corresponding to the no-damage SN curve under the same load level are obtained. Divide the fatigue life corresponding to the damage SN curve by the fatigue life corresponding to the no-damage SN curve to obtain the damage low-cycle fatigue reduction factor for the corresponding damage depth. Based on the damage low-cycle fatigue reduction coefficient corresponding to different damage depths, a correspondence between damage depth and damage low-cycle fatigue reduction coefficient is constructed.

9. The method of claim 8, wherein, The target aerospace part and the aerospace part to be evaluated are made of the same type of material, and the method further includes: When the damage SN curve corresponding to the target aerospace part is not established, the crack initiation life of the target aerospace part can be directly estimated based on the correspondence between the damage depth and the damage low-cycle fatigue reduction coefficient and the undamaged SN curve corresponding to the target aerospace part.

10. A low-cycle fatigue life assessment system for aircraft parts, characterized in that, The method for assessing the low-cycle fatigue life of aircraft parts for damage according to any one of claims 1 to 9 includes: The first acquisition module is used to acquire the current damage depth and working stress of the aerospace part to be evaluated; The second acquisition module is used to acquire a set of SN curves of a sample material that is the same material as the aerospace part to be evaluated, which is pre-established, wherein the set of SN curves includes the damage SN curves of the sample material at at least two different damage depths. The first acquisition module is used to obtain the crack initiation life by interpolating the set of SN curves based on the current damage depth and the working stress. The processing module is used to take the sum of the current damage depth and the preset correction value as the initial crack size, perform crack propagation analysis based on the initial crack size to obtain the initial crack propagation life, and perform conservative processing on the initial crack propagation life to obtain the target crack propagation life, wherein the conservative processing is used to represent the reduction processing of the initial crack propagation life. The second acquisition module is used to sum the crack initiation life and the target crack propagation life to obtain the total low-cycle fatigue life of the aerospace part to be evaluated.

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