Residual life prediction method applied to aero-engine hot end component

Abstract

The invention discloses a residual life prediction method applied to a hot end component of an aero-engine, which comprises the following steps of: performing finite element analysis on the hot end component of the aero-engine with cracks, and determining J integral and stress three-axis degree of a crack front edge node under different loads; the representative position of the crack front edge is determined by combining the J integral of the crack front edge and the stress three-axis degree processing result; generating a traditional failure assessment diagram through the J integral of the representative position of the crack front edge; a correlation function of crack driving force and ligament yield parameters under different crack sizes is constructed, and a failure assessment diagram and a damage tolerance failure assessment diagram combined with the crack size effect are generated; and performing allowable crack size and residual fatigue life prediction to obtain a fatigue life prediction result. According to the method, the residual life prediction precision and reliability of the aero-engine hot end component are improved, the service cycle and maintenance strategy of equipment are optimized, and potential safety hazards caused by structural failure are reduced.

Description

Technical Field

[0001] The invention relates to the field of aeroengines, and more particularly to a method for predicting the remaining life of a hot end component of an aeroengine. Background Art

[0002] Aircraft engine hot end components, such as aircraft engine turbine disks, operate under extreme conditions and often face structural integrity problems caused by cracks. These cracks not only directly threaten the strength and durability of the components, but may also pose a major risk to the safe operation of the entire equipment. Traditional remaining life prediction methods have obvious limitations when facing cracked aircraft engine hot end components.

[0003] At present, the research on the prediction of the remaining fatigue life of the hot end components of aircraft engines based on the allowable crack size has a mature method flow, mainly including empirical methods, finite element methods and fracture mechanics methods. Although the empirical method is simple to operate, the results are too conservative and rely on a large amount of experimental data and engineering experience, which makes it difficult to meet the requirements of high-precision prediction of the hot end components of modern aircraft engines. Although the finite element method can provide detailed stress distribution information, the calculation cost is high, professional software and complex constitutive models are required, and the required parameters are difficult to obtain. Although the fracture mechanics method is simple to operate, it is usually based on the assumptions of quasi-static fracture and plane strain state, which is inconsistent with the actual crack extension under fatigue load, resulting in poor prediction results.

[0004] In addition, the traditional failure assessment diagram (FAD) method is mainly used to evaluate the crack propagation behavior of materials, but when dealing with cracked structures, it often ignores the influence of factors such as crack size, structural geometry, and material nonlinear behavior, resulting in overly conservative or inaccurate prediction results. For example, the traditional FAD method is usually based on the plane strain state of the crack tip of the standard specimen, while the actual stress state of the crack tip of the hot end component of the aircraft engine is more complex, which further affects the accuracy of the prediction.

[0005] Therefore, how to design a remaining life prediction method for aircraft engine hot end components that can comprehensively consider crack size effects, load conditions and material properties, and improve the remaining life prediction accuracy and reliability of aircraft engine hot end components is an urgent problem that technical personnel in this field need to solve. Summary of the invention

[0006] In view of this, the present invention provides a remaining life prediction method applied to the hot end components of aircraft engines. By combining finite element analysis and fracture mechanics theory, the J-integral and stress triaxiality of the crack front are accurately obtained, and a new failure assessment diagram considering the crack size effect is generated, which significantly improves the accuracy and reliability of the prediction of the allowable crack size and remaining fatigue life of the hot end components of aircraft engines.

[0007] In order to achieve the above object, the present invention adopts the following technical solution:

[0008] A method for predicting the remaining life of a hot end component of an aero-engine comprises the following steps:

[0009] S1. Perform finite element analysis on the hot end parts of the aero-engine with cracks to determine the J integral and stress triaxiality of the crack front node under different loads;

[0010] S2. Extrapolating the stress triaxiality of the crack front node to obtain the stress triaxiality extrapolation result of the crack front node;

[0011] S3, based on the J-integral of the crack front node and the stress triaxiality extrapolation processing results, combined with the preset threshold, determine the representative position of the crack front;

[0012] S4, generating a traditional failure assessment diagram through the J-integral of the representative position of the crack front;

[0013] S5. Using the corresponding coordinates on the traditional failure assessment diagram, constructing a correlation function between the crack driving force and the ligament yield parameter under different crack sizes;

[0014] S6. Generate a failure assessment diagram combining crack size effect through the correlation function between crack driving force and ligament yield parameter under different crack sizes and loads;

[0015] S7, generating a damage tolerance failure assessment diagram based on the failure assessment diagram combined with the crack size effect;

[0016] S8. Based on the damage tolerance failure assessment diagram, predict the allowable crack size to obtain an allowable crack size prediction result;

[0017] S9. Based on the prediction result of the allowable crack size and combined with the crack propagation model, the remaining fatigue life is predicted to obtain the fatigue life prediction result.

[0018] Furthermore, the S1 includes:

[0019] Obtain material parameters of the materials constituting the hot end components of aircraft engines;

[0020] Based on the finite element model, the singularity of the mesh near the crack front is set, and different external loads are applied;

[0021] The material parameters are input into the finite element model to obtain the J integral of the crack front, and the principal stress is extracted to obtain the stress triaxiality of the crack front. The J integral is obtained through the J integral distribution diagram at different angles between the crack front node and the crack width, including the elastic J integral and the elastoplastic J integral.

[0022] Furthermore, in S2, the stress triaxiality of the crack front node is determined by different The stress triaxiality distribution diagram under the ratio is obtained; where r represents the distance from the node to the crack front, σ Y represents yield stress, and J represents J-integral.

[0023] Furthermore, in S3, the representative positions of the crack front include: nodes where the J integral is in the range of 0.8 to 1 times the maximum J integral and the stress triaxiality is in the range of 0.9 to 1 times the maximum value.

[0024] Furthermore, in S4, the traditional failure assessment diagram is expressed as:

[0025]

[0026] Among them, f(L r ) represents the crack driving force and is the ordinate of the failure assessment curve; L r is the ligament yield parameter, is the abscissa of the failure assessment curve; J e represents the elastic J integral, J ec represents the elastic-plastic J integral, σ ref represents the reference stress, σ 0.2 represents the material yield strength, σ app represents the remote applied stress, σ0 represents the reference yield stress, represents the critical ligament yield parameter, σ u Indicates the tensile strength of the material;

[0027] And when L r =1,σ0=σ app ,satisfy:

[0028]

[0029]

[0030] Where E represents Young's modulus, ε ref represents the reference strain, J e (σ app ), J ec(σapp) are the far-end applied stress σ app Elastic J integral and elastoplastic J integral under .

[0031] Furthermore, the S5 includes:

[0032] Based on the J integral and the crack size a, crack driving force f(L r ), through the crack size a on the crack driving force f(L r ) to represent:

[0033] f(L r (a))=P+Ma+Na 2 +Ha 3

[0034] Among them, f(L r (a)) represents the crack driving force related to the crack size, and P, M, N, and H represent the fitting parameters;

[0035] Based on the correlation between the reference yield stress σ0 and the crack size a and J integral, the ligament yield parameter Lr is expressed by the crack size a:

[0036]

[0037] σ0(a)=P′+M′a+N′a 2

[0038] And when there is no crack in the component, σ0(a=0)=σ 0.2 ,satisfy:

[0039] σ0(a)=σ 0.2 +M′a+N′a 2

[0040] Among them, L r (a) represents the ligament yield parameter related to the crack size, σ0(a) represents the reference yield stress related to the crack size, and P′, M′, and N′ represent the fitting parameters.

[0041] Further, in the failure assessment diagram combined with the crack size effect, the vertical coordinate of the failure assessment curve is the crack driving force f(L r (a)), the abscissa of the failure curve is the ligament yield parameter L related to the crack size r (a);

[0042] Critical ligament yield parameter It is expressed as:

[0043]

[0044] Among them, the critical ligament yield parameter represents the critical ligament yield parameter enlarge times.

[0045] Furthermore, in S7, the damage tolerance failure assessment diagram includes: a meaningless area, a non-crack extension area, a safe area, and a dangerous area.

[0046] Further, in said S8, the allowable crack size is determined by the coordinates of the intersection of the damage tolerance failure assessment curve and the assessment curve;

[0047] The evaluation curve calculation is expressed as:

[0048]

[0049] Among them, K r represents the ordinate of the evaluation curve, K(a) represents the stress intensity factor of different crack sizes, K mat represents fracture toughness, L r (a) represents the horizontal axis of the evaluation curve.

[0050] Furthermore, in S9, the remaining fatigue life is expressed as:

[0051]

[0052] Among them, N residual represents the remaining fatigue life, a0 represents the initial crack size, a cr represents the allowable crack size, ΔK represents the stress intensity factor range, area represents the projection area of ​​the crack in the loading direction, and c and n represent material constants.

[0053] It can be seen from the above technical solution that compared with the prior art, the technical solution of the present invention has the following advantages:

[0054] Beneficial effects:

[0055] 1. This method accurately calculates the J-integral and stress triaxiality of the crack front by performing finite element analysis on the hot end components of aircraft engines with different crack sizes, and generates a damage tolerance failure assessment diagram that takes into account the crack size effect based on these data. The influence of crack size on crack propagation behavior is considered, and the prediction accuracy is improved by extrapolating the J-integral and stress triaxiality.

[0056] 2. When constructing the correlation function between crack driving force and ligament yield parameter under different crack sizes, the data basis of the traditional failure assessment diagram was used, and an improved failure assessment diagram was proposed in combination with the crack size effect. This enhances the flexibility of the method and makes it applicable to a wider range of crack sizes and load conditions.

[0057] 3. Taking into account various factors such as crack size, load conditions and material properties, the allowable crack size is predicted based on the damage tolerance failure assessment diagram, and the remaining fatigue life is predicted in combination with the crack growth model. The safety performance of the hot end components of aircraft engines can be evaluated more scientifically, and reasonable maintenance and replacement plans can be formulated to ensure the safe operation of the equipment and extend its service life. BRIEF DESCRIPTION OF THE DRAWINGS

[0058] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are only embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on the provided drawings without paying creative work.

[0059] Figure 1 A flow chart of a method for predicting the remaining life of a hot end component of an aero-engine provided in an embodiment of the present invention;

[0060] Figure 2 A J-integral distribution diagram of a crack front provided by an embodiment of the present invention;

[0061] Figure 3 A stress triaxiality distribution diagram of a crack front provided by an embodiment of the present invention;

[0062] Figure 4 A damage tolerance failure assessment diagram provided by an embodiment of the present invention;

[0063] Figure 5 A schematic diagram of prediction results of allowable crack size provided by an embodiment of the present invention;

[0064] Figure 6 A schematic diagram of the remaining fatigue life prediction results provided by an embodiment of the present invention. DETAILED DESCRIPTION

[0065] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0066] like Figure 1 As shown, this embodiment provides a method for predicting the remaining life of a hot end component of an aircraft engine, comprising the following steps:

[0067] S1. Perform finite element analysis on the hot end parts of the aero-engine with cracks to determine the J integral and stress triaxiality of the crack front node under different loads;

[0068] S2. Extrapolating the stress triaxiality of the crack front node to obtain the stress triaxiality extrapolation result of the crack front node;

[0069] S3, based on the J-integral of the crack front node and the stress triaxiality extrapolation processing results, combined with the preset threshold, determine the representative position of the crack front;

[0070] S4, generating a traditional failure assessment diagram through the J-integral of the representative position of the crack front;

[0071] S5. Using the corresponding coordinates on the traditional failure assessment diagram, constructing a correlation function between the crack driving force and the ligament yield parameter under different crack sizes;

[0072] S6. Generate a failure assessment diagram combining crack size effect through the correlation function between crack driving force and ligament yield parameter under different crack sizes and loads;

[0073] S7, generating a damage tolerance failure assessment diagram based on the failure assessment diagram combined with the crack size effect;

[0074] S8. Based on the damage tolerance failure assessment diagram, predict the allowable crack size to obtain an allowable crack size prediction result;

[0075] S9. Based on the prediction result of the allowable crack size and combined with the crack propagation model, the remaining fatigue life is predicted to obtain the fatigue life prediction result.

[0076] The remaining life prediction method here combines finite element analysis, failure assessment diagrams and crack propagation models. It can comprehensively consider crack size effects, load conditions and material properties, and achieve high-precision prediction of the allowable crack size and remaining fatigue life of aircraft engine hot end components. It provides a scientific and reliable basis for related engineering design and maintenance, effectively reduces the risk of structural failure, and improves the reliability and safety of equipment.

[0077] Powder high temperature alloy FGH96 material is often used as the constituent material of the hot end components of aircraft engines. The following takes a semicircular artificial crack specimen made of powder high temperature alloy FGH96 material as an example to further explain in detail the various steps and related technical features in the above-mentioned remaining life prediction method:

[0078] The material parameters of powder high temperature alloy FGH96 are shown in Tables 1 and 2 below:

[0079] Table 1

[0080]

[0081] Table 2

[0082]

[0083] The fatigue test data of powder high temperature alloy FGH96 material under stress ratio of 0.05 are shown in Table 3 below:

[0084] Table 3

[0085]

[0086]

[0087] In this embodiment S1, a finite element analysis is performed on a hot end component of an aircraft engine with cracks to determine the J integral and stress triaxiality of multiple nodes at the crack front under different loads;

[0088] Specifically include:

[0089] Obtain the material parameters of the material constituting the hot end component of the aircraft engine; the material parameters here are the parameters in Table 1 above; based on the finite element model, set the singularity of the grid near the crack front, and apply different external loads; input the material parameters into the finite element model, obtain the J integral of the crack front, and extract the principal stress to obtain the stress triaxiality of the crack front;

[0090] like Figure 2 As shown in the figure, the J integral is a parameter that measures the energy release rate at the crack tip. It is obtained through the J integral distribution diagram at different angles between the crack front node and the crack width, including the elastic J integral and the elastic-plastic J integral. The elastic J integral mainly reflects the elastic deformation energy at the crack tip, while the elastic-plastic J integral takes into account the influence of plastic deformation.

[0091] In this figure, the horizontal axis represents the angle between the crack front node and the crack width. The ordinate represents the J-integral. There are four curves, each representing a typical combination of crack length (a) and stress level (b). Specifically, they include: blue solid line: a=1200μm, b=900MPa; blue dashed line: a=1200μm, b=1200MPa; red solid line: a=1800μm, b=900MPa; red dashed line: a=1800μm, b=1000MPa; In addition, each curve is marked with the corresponding maximum value of the J-integral (max), as well as the J-integral value when the angle between the crack front node and the crack width is the largest;

[0092] The specific values ​​are as follows: The blue solid line, that is, when a = 1200μm and b = 900MPa, the maximum value of J integral is 12.8N / mm. The maximum J integral is 10.5N / mm; the blue dotted line, that is, when a=1200μm and b=1200MPa, the maximum J integral is 109.4N / mm. The maximum J integral is 109.0N / mm; the red solid line, that is, when a=1800μm, b=900MPa, the maximum J integral is 38.2N / mm. The maximum J integral is 33.0N / mm; the red dotted line, that is, when a=1800μm and b=1000MPa, the maximum J integral is 76.3N / mm. The J integral at maximum is 74.4N / mm;

[0093] From this figure, when the crack length is the same, a higher stress level will result in a higher J-integral value. When the stress level is the same, a longer crack length will also result in a higher J-integral value. As the value of increases, the J integral value gradually rises to a maximum value, and then tends to be stable or slightly decreases. This shows that there is an optimal angle that makes the J integral reach its maximum value.

[0094] In addition, in finite element analysis, it is also necessary to extract the principal stress values ​​of multiple nodes at the crack front. According to the principal stress values, the stress triaxiality of each node is calculated using the definition and calculation formula of stress triaxiality. Stress triaxiality is an important parameter that reflects the tendency of material failure under multiaxial stress state.

[0095] Stress triaxiality calculation formula:

[0096]

[0097] Among them, σ h is the hydrostatic stress, σ Mises is the von Mises stress, σ1, σ2, and σ3 are principal stresses, representing the maximum tensile stress, intermediate stress, and minimum compressive stress, respectively.

[0098] In this embodiment S2, the stress triaxiality of the crack front node is extrapolated to obtain the stress triaxiality extrapolation result of the crack front node;

[0099] like Figure 3 As shown in the figure, the stress triaxiality of the crack front node is The stress triaxiality distribution diagram under the ratio is obtained; where r represents the distance from the node to the crack front, σ Y represents yield stress, and J represents J-integral.

[0100] In this figure, the horizontal axis represents ratio, r represents the distance from the node to the crack front, σ Y represents yield stress, J represents J integral; the ordinate represents stress triaxiality The black dots in the figure represent the data points actually measured or calculated, and the red oblique line is a fitting curve obtained by extrapolating these data points, which clarifies the stress triaxiality h with Increasing and gradually decreasing trend.

[0101] And when When it approaches zero, the stress triaxiality The initial value of is about 2.78, which is the maximum stress triaxiality; as the distance r from the node to the crack front or As the ratio increases, the stress triaxiality h will show a downward trend, and the stress state of the material becomes more uniform or closer to the uniaxial tension state at a position far away from the crack tip.

[0102] When performing the extrapolation of stress triaxiality, the actual stress triaxiality of several nodes at the crack front is first obtained by finite element analysis. Then, several typical nodes near the crack front are selected, and a straight line is constructed through these known points based on the trend of their stress triaxiality changing with distance. This straight line can assist in predicting the stress triaxiality of nodes beyond the known range. That is, the linear relationship between the existing data points is used to estimate the stress triaxiality of the unknown position, so as to achieve effective extrapolation of the stress triaxiality of the external nodes.

[0103] In this embodiment S3, based on the J-integral of the crack front node and the stress triaxiality extrapolation processing result, the representative position of the crack front is determined in combination with a preset threshold;

[0104] Specifically, in the finite element analysis, the J integral and stress triaxiality extrapolation results of the crack front nodes have been calculated. In order to determine the representative position of the crack front, it is first necessary to find the maximum value of the J integral and the maximum value of the stress triaxiality extrapolation results in the node. Then, the nodes whose J integral is within the range of 0.8 to 1 times the maximum J integral and the stress triaxiality extrapolation results are within the range of 0.9 to 1 times the maximum value are selected. These nodes can better reflect the crack propagation driving force of the crack front and reflect the local stress state of the crack front.

[0105] Furthermore, the above-selected nodes are cross-screened, and the nodes that meet the representative conditions of both the J integral and the stress triaxiality extrapolation results are selected as the final representative positions, and the nodes with the deepest cracks are selected as representative positions. Finally, the selected representative positions are data verified to ensure the rationality and consistency of their J integral and stress triaxiality extrapolation results, and the selection of representative positions is verified to be reasonable by comparing with the data of other nodes.

[0106] In this embodiment S4, a conventional failure assessment diagram is generated by J-integral of a representative position of the crack front;

[0107] The traditional failure assessment diagram is expressed as:

[0108]

[0109] Among them, f(L r ) represents the crack driving force and is the ordinate of the failure assessment curve; L r is the ligament yield parameter, is the abscissa of the failure assessment curve; J e represents the elastic J integral, J ecrepresents the elastic-plastic J integral, σ ref represents the reference stress, σ 0.2 represents the material yield strength, σ app represents the remote applied stress, σ0 represents the reference yield stress, represents the critical ligament yield parameter, σ u Indicates the tensile strength of the material;

[0110] And when L r =1,σ0=σ app ,satisfy:

[0111]

[0112] Where E represents Young's modulus, ε ref represents the reference strain, J e (σ app ), J ec(σapp) are the far-end applied stress σ app Elastic J integral and elastoplastic J integral under .

[0113] In this embodiment S5, the corresponding coordinates on the traditional failure assessment diagram are used to construct a correlation function between the crack driving force and the ligament yield parameter under different crack sizes;

[0114] Specifically include:

[0115] Based on the J integral and the crack size a, crack driving force f(L r ), through the crack size a on the crack driving force f(L r ) to represent:

[0116] f(L r (a))=P+Ma+Na 2 +Ha 3

[0117] Among them, f(L r (a)) represents the crack driving force related to the crack size, and P, M, N, and H represent the fitting parameters;

[0118] Based on the correlation between the reference yield stress σ0 and the crack size a and J integral, the ligament yield parameter Lr is expressed by the crack size a:

[0119]

[0120] σ0(a)=P'+M'a+N'a 2

[0121] And when there is no crack in the component, σ0(a=0)=σ 0.2 ,satisfy:

[0122] σ0(a)=σ 0.2 +M′a+N′a 2

[0123] Among them, L r (a) represents the ligament yield parameter related to the crack size, σ0(a) represents the reference yield stress related to the crack size, and P′, M′, and N′ represent the fitting parameters.

[0124] Furthermore, through the correlation function between the crack driving force and the ligament yield parameter under different crack sizes and loads, a failure assessment diagram combining the crack size effect is generated;

[0125] In the failure assessment diagram combining the crack size effect, the vertical coordinate of the failure assessment curve is the crack driving force f(L) related to the crack size. r (a)), the abscissa of the failure curve is the ligament yield parameter L related to the crack size r (a);

[0126] Critical ligament yield parameter It is expressed as:

[0127]

[0128] Among them, the critical ligament yield parameter represents the critical ligament yield parameter enlarge times.

[0129] like Figure 4 As shown, a damage tolerance failure assessment diagram is generated based on the failure assessment diagram combined with the crack size effect;

[0130] The damage tolerance failure assessment diagram includes: meaningless area, no crack extension area, safe area and dangerous area.

[0131] Specifically, the meaningless area is Area; The area without crack extension is Area, where K th is the fatigue crack extension threshold; the safety zone is and The enclosed area; the danger zone is or f(L r (a))>K r area.

[0132] Further, based on the damage tolerance failure assessment diagram, an allowable crack size prediction is performed to obtain an allowable crack size prediction result;

[0133] The allowable crack size is determined by the coordinates of the intersection of the damage tolerance failure assessment curve and the assessment curve;

[0134] The evaluation curve calculation is expressed as:

[0135]

[0136] Among them, K r represents the ordinate of the evaluation curve, K(a) represents the stress intensity factor of different crack sizes, K mat represents fracture toughness, L r (a) represents the horizontal axis of the evaluation curve.

[0137] Furthermore, based on the prediction results of the allowable crack size, the remaining fatigue life is predicted in combination with the crack growth model to obtain the fatigue life prediction results;

[0138] The remaining fatigue life is expressed as:

[0139]

[0140] Among them, N residual represents the remaining fatigue life, a0 represents the initial crack size, a cr represents the allowable crack size, ΔK represents the stress intensity factor range, area represents the projection area of ​​the crack in the loading direction, and c and n represent material constants.

[0141] like Figure 5 As shown, the comparison results of the allowable crack size and the actual size of the powder high-temperature alloy FGH96 material predicted by the remaining life prediction method of this embodiment have a prediction accuracy within ±1.2 times the dispersion band and an error within ±20%.

[0142] like Figure 6 As shown, the comparison results of the remaining fatigue life of the powder high-temperature alloy FGH96 material predicted by the remaining life prediction method of this embodiment and the actual remaining life, the prediction accuracy of different initial crack sizes and different temperature conditions is within the ±2 times dispersion band.

[0143] The remaining life prediction method provided in this embodiment is aimed at the hot end components of aircraft engines, such as aircraft engine turbine disks, and accurately predicts the crack size and remaining fatigue life by comprehensively applying finite element analysis, failure assessment diagrams and crack growth models. It fully considers the crack size effect, load conditions and material properties, significantly improves the prediction accuracy, and provides strong technical support for the damage tolerance design and in-service maintenance strategy of the hot end components of aircraft engines, which can effectively reduce the risk of structural failure and improve the reliability and safety of equipment.

[0144] In this specification, each embodiment is described in a progressive manner, and each embodiment focuses on the differences from other embodiments. The same or similar parts between the embodiments can be referred to each other. For the system disclosed in the embodiment, since it corresponds to the method disclosed in the embodiment, the description is relatively simple, and the relevant parts can be referred to the method part.

[0145] The above description of the disclosed embodiments enables one skilled in the art to implement or use the present invention. Various modifications to these embodiments will be apparent to one skilled in the art, and the general principles defined herein may be implemented in other embodiments without departing from the spirit or scope of the present invention. Therefore, the present invention will not be limited to the embodiments shown herein, but rather to the widest scope consistent with the principles and novel features disclosed herein.

Claims

1. A method for predicting the remaining life of a hot end component of an aero-engine, characterized in that: The following steps are involved: S1. Perform finite element analysis on the hot end parts of the aero-engine with cracks to determine the J integral and stress triaxiality of the crack front node under different loads; S2. Extrapolating the stress triaxiality of the crack front node to obtain the stress triaxiality extrapolation result of the crack front node; S3, based on the J-integral of the crack front node and the stress triaxiality extrapolation processing results, combined with the preset threshold, determine the representative position of the crack front; S4, generating a traditional failure assessment diagram through the J-integral of the representative position of the crack front; S5. Using the corresponding coordinates on the traditional failure assessment diagram, constructing a correlation function between the crack driving force and the ligament yield parameter under different crack sizes; S6. Generate a failure assessment diagram combining crack size effect through the correlation function between crack driving force and ligament yield parameter under different crack sizes and loads; S7, generating a damage tolerance failure assessment diagram based on the failure assessment diagram combined with the crack size effect; S8. Based on the damage tolerance failure assessment diagram, predict the allowable crack size to obtain an allowable crack size prediction result; S9. Based on the prediction result of the allowable crack size and combined with the crack propagation model, the remaining fatigue life is predicted to obtain the fatigue life prediction result.

2. The method for predicting the remaining life of a hot end component of an aero-engine according to claim 1, characterized in that: Said S1 comprises: Obtain material parameters of the materials constituting the hot end components of aircraft engines; Based on the finite element model, the singularity of the mesh near the crack front is set, and different external loads are applied; The material parameters are input into the finite element model to obtain the J integral of the crack front, and the principal stress is extracted to obtain the stress triaxiality of the crack front. The J integral is obtained through the J integral distribution diagram at different angles between the crack front node and the crack width, including the elastic J integral and the elastoplastic J integral.

3. The method for predicting the remaining life of a hot end component of an aero-engine according to claim 1, characterized in that: In S2, the stress triaxiality of the crack front node is determined by different The stress triaxiality distribution diagram under the ratio is obtained; where r represents the distance from the node to the crack front, σ Y represents yield stress, and J represents J-integral.

4. The method for predicting the remaining life of a hot end component of an aero-engine according to claim 1, characterized in that: In S3, the representative positions of the crack front include nodes where the J integral is within the range of 0.8 to 1 times the maximum J integral and the stress triaxiality is within the range of 0.9 to 1 times the maximum value.

5. The method for predicting the remaining life of a hot end component of an aero-engine according to claim 1, characterized in that: In S4, the traditional failure assessment diagram is expressed as: Among them, f(L r ) represents the crack driving force and is the ordinate of the failure assessment curve; L r is the ligament yield parameter, is the abscissa of the failure assessment curve; J e represents the elastic J integral, J ec represents the elastic-plastic J integral, σ ref represents the reference stress, σ 0.2 represents the material yield strength, σ app represents the remote applied stress, σ0 represents the reference yield stress, represents the critical ligament yield parameter, σ u Indicates the tensile strength of the material; And when L r =1,σ0=σ app ,satisfy: Where E represents Young's modulus, ε ref represents the reference strain, J e (σ app ), J ec(σapp) are the far-end applied stress σ app Elastic J integral and elastoplastic J integral under .

6. The method for predicting the remaining life of a hot end component of an aero-engine according to claim 1, characterized in that: The S5 comprises: Based on the J integral and the crack size a, crack driving force f(L r ), through the crack size a on the crack driving force f(L r ) to represent: f(L r (a))=P+Ma+Na 2 + Ha 3 Among them, f(L r (a)) represents the crack driving force related to the crack size, and P, M, N, and H represent the fitting parameters; Based on the correlation between the reference yield stress σ0 and the crack size a and J integral, the ligament yield parameter Lr is expressed by the crack size a: σ0(a)=P'+M′a+N'a 2 And when there is no crack in the component, σ0(a=0)=σ 0.2 ,satisfy: σ0(a)=σ 0.2 +M′a+N'a 2 Among them, L r (a) represents the ligament yield parameter related to the crack size, σ0(a) represents the reference yield stress related to the crack size, and P′, M′, and N′ represent the fitting parameters.

7. The method for predicting the remaining life of a hot end component of an aero-engine according to claim 1, characterized in that: In the failure assessment diagram combined with the crack size effect, the vertical coordinate of the failure assessment curve is the crack driving force f(L) related to the crack size. r (a)), the abscissa of the failure curve is the ligament yield parameter L related to the crack size r (a); Critical ligament yield parameter It is expressed as: Among them, the critical ligament yield parameter represents the critical ligament yield parameter enlarge times.

8. The method for predicting the remaining life of a hot end component of an aero-engine according to claim 1, characterized in that: In S7, the damage tolerance failure assessment diagram includes: a meaningless area, a non-crack extension area, a safe area, and a dangerous area.

9. The method for predicting the remaining life of a hot end component of an aero-engine according to claim 1, characterized in that: In said S8, the allowable crack size is determined by the coordinates of the intersection of the damage tolerance failure assessment curve and the assessment curve; The evaluation curve calculation is expressed as: Among them, K r represents the ordinate of the evaluation curve, K(a) represents the stress intensity factor of different crack sizes, K mat represents fracture toughness, L r (a) represents the horizontal axis of the evaluation curve.

10. The method for predicting the remaining life of a hot end component of an aero-engine according to claim 1, characterized in that: In S9, the remaining fatigue life is expressed as: Among them, N residual represents the remaining fatigue life, a0 represents the initial crack size, a cr represents the allowable crack size, ΔK represents the stress intensity factor range, area represents the projection area of ​​the crack in the loading direction, and c and n represent material constants.

Images