A method for predicting the remaining life of hot-end components of aero-engines

Abstract

The present invention discloses a method for predicting the remaining life of hot-end components of an aero-engine, comprising the following steps: performing finite element analysis on the hot-end components of the aero-engine with cracks to determine the J-integral and stress triaxiality of the crack front nodes under different loads; combining the processing results of the J-integral and stress triaxiality of the crack front to determine the representative position of the crack front; generating a traditional failure assessment diagram through the J-integral of the representative position of the crack front; and constructing a correlation function between the crack driving force and the ligament yield parameter under different crack sizes to generate a failure assessment diagram considering crack size effect and a damage tolerance failure assessment diagram; performing prediction of the allowable crack size and remaining fatigue life to obtain the fatigue life prediction result. This method improves the prediction accuracy and reliability of the remaining life of hot-end components of an aero-engine, helps to optimize the service cycle and maintenance strategy of the equipment, and reduces the safety hazards caused by structural failure.

Description

Technical Field

[0001] The present invention relates to the field of aero-engines, and more particularly to a method for predicting the remaining life of hot-end components of aero-engines. Background Art

[0002] Hot-end components of aero-engines, such as turbine disks of aero-engines, 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 also may pose significant risks to the safe operation of the entire equipment. Traditional methods for predicting the remaining life have obvious limitations when dealing with hot-end components of aero-engines containing cracks.

[0003] Currently, research on predicting the remaining fatigue life of hot-end components of aero-engines around the allowable crack size has a mature method process, 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, making it difficult to meet the requirements of high-precision prediction for modern hot-end components of aero-engines. Although the finite element method can provide detailed stress distribution information, the calculation cost is high, requiring professional software and complex constitutive models, and it is difficult to obtain the required parameters. 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 does not conform to the crack propagation under actual fatigue loads, resulting in poor accuracy of the 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 at the crack tip of a standard specimen, while the stress state at the crack tip of actual hot-end components of aero-engines is more complex, which further affects the accuracy of the prediction.

[0005] Therefore, how to design a method for predicting the remaining life of hot-end components of aero-engines that can comprehensively consider the crack size effect, load conditions, and material properties, and improve the accuracy and reliability of predicting the allowable crack size and remaining fatigue life of hot-end components of aero-engines is an urgent problem to be solved by those skilled in the art. Summary of the Invention

[0006] In view of this, the present invention provides a method for predicting the remaining life of hot-end components of aero-engines. By combining finite element analysis and fracture mechanics theory, the J-integral and stress triaxiality at 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 predicting the allowable crack size and remaining fatigue life of hot-end components of aero-engines.

[0007] To achieve the above object, the present invention adopts the following technical solutions:

[0008] A method for predicting the remaining life of hot-end components of an aeroengine, comprising the following steps:

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

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

[0011] S3. Based on the J integral of the crack front nodes and the extrapolation result of the stress triaxiality, determine the representative position of the crack front in combination with a preset threshold;

[0012] S4. Generate a traditional failure assessment diagram through the J integral at the representative position of the crack front;

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

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

[0015] S7. Generate a damage tolerance failure assessment diagram based on the failure assessment diagram considering the crack size effect;

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

[0017] S9. Based on the prediction result of the allowable crack size, combine with the crack propagation model to perform prediction of the remaining fatigue life to obtain the prediction result of the fatigue life.

[0018] Further, the S1 includes:

[0019] Obtain the material parameters of the materials constituting the hot-end components of the aeroengine;

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

[0021] Input the material parameters into the finite element model to obtain the J integral of the crack front, and perform principal stress extraction to obtain the stress triaxiality of the crack front; wherein, the J integral is obtained through the J integral distribution diagram at different angles between the crack front nodes and the crack width, including the elastic J integral and the elastic-plastic J integral.

[0022] Further, in S2, the stress triaxiality of the crack front nodes is obtained from the stress triaxiality distribution diagrams at different ratios; where r represents the distance from the node to the crack front, σ Y represents the yield stress, and J represents the J-integral.

[0023] Further, in S3, the representative positions of the crack front include: the 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.

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

[0025]

[0026] where f(L r ) represents the crack driving force, which is the ordinate of the failure assessment curve; L r is the ligament yield parameter, which 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 remotely applied stress, σ0 represents the reference yield stress, represents the critical ligament yield parameter, σ u represents the material tensile strength;

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

[0028]

[0029]

[0030] where E represents Young's modulus, ε ref represents the reference strain, J e (σ app ), J ec(σapp) respectively represent the elastic J-integral and the elastic-plastic J-integral under the remotely applied stress σ app .

[0031] Further, S5 includes:

[0032] Based on the correlation between the J-integral and the crack size a and the crack driving force f(L r ), the crack driving force f(L r ) is expressed in terms of the crack size a:

[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 fitting parameters;

[0035] Based on the correlation relationship between the reference yield stress σ0, the crack size a, and the 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 the component has no crack, σ0(a = 0) = σ 0.2 , satisfying:

[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′, N′ represent fitting parameters.

[0041] Furthermore, in the S6, in the failure assessment diagram considering the crack size effect, the ordinate of the failure assessment curve is the crack driving force f(L r (a)) related to the crack size, and the abscissa of the failure curve is the ligament yield parameter L r (a);

[0042] Critical ligament yield parameter is expressed as:

[0043]

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

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

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

[0047] The calculation of the assessment curve is expressed as:

[0048]

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

[0050] Further, 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 projected area of the crack in the loading direction, and c, n represent material constants.

[0053] Through the above technical solutions, compared with the prior art, the technical solutions of the present invention have the following

[0054] beneficial effects:

[0055] 1. By performing finite element analysis on the hot-end components of aero-engines with different crack sizes, the J-integral and stress triaxiality at the crack front are accurately calculated, and a damage tolerance failure assessment diagram considering the crack size effect is generated based on these data. The influence of crack size on crack propagation behavior is considered, and the prediction accuracy is improved through the extrapolation processing of the J-integral and stress triaxiality.

[0056] 2. When constructing the correlation function between the crack driving force and the ligament yield parameter under different crack sizes, the data basis of the traditional failure assessment diagram is utilized, and combined with the crack size effect, an improved failure assessment diagram is proposed. The flexibility of the method is enhanced, and it can be applied to a wider range of crack sizes and load conditions.

[0057] 3. Various factors such as crack size, load conditions, and material properties are comprehensively considered. 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 propagation model. The safety performance of the hot-end components of aero-engines can be evaluated more scientifically, and reasonable maintenance and replacement plans can be formulated, thereby ensuring the safe operation of the equipment and extending its service life. Brief Description of the Drawings

[0058] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the accompanying drawings required for the description of the embodiments or the prior art. Obviously, the accompanying drawings in the following description are only the embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other accompanying drawings can also be obtained based on the provided accompanying drawings.

[0059] Figure 1 It is a flowchart of a remaining life prediction method applied to the hot-end components of an aero-engine provided by an embodiment of the present invention;

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

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

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

[0063] Figure 5 It is a schematic diagram of the predicted result of the allowable crack size provided by an embodiment of the present invention;

[0064] Figure 6 It is a schematic diagram of the predicted result of the remaining fatigue life provided by an embodiment of the present invention. Detailed implementation manners

[0065] The following will clearly and completely describe the technical solutions in the embodiments of the present invention with reference to the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts belong to the scope of protection of the present invention.

[0066] As Figure 1 shown, this embodiment provides a remaining life prediction method applied to the hot-end components of an aero-engine, including the following steps:

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

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

[0069] S3. Based on the J integral of the crack front nodes and the extrapolation result of the stress triaxiality, determine the representative position of the crack front in combination with a preset threshold value;

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

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

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

[0073] S7. Generate a damage tolerance failure assessment diagram based on the failure assessment diagram considering the crack size effect;

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

[0075] S9. Based on the prediction result of the allowable crack size, combine the crack propagation model to perform the prediction of the remaining fatigue life to obtain the prediction result of the fatigue life.

[0076] The remaining life prediction method herein combines finite element analysis, failure assessment diagrams, and crack propagation models, and can comprehensively consider the crack size effect, load conditions, and material properties, achieving high-precision prediction of the allowable crack size and remaining fatigue life of the hot-end components of aero-engines, providing a scientific and reliable basis for relevant engineering designs and maintenance, effectively reducing the risk of structural failure, and improving the reliability and safety of the equipment.

[0077] The powder superalloy FGH96 material is often used as the constituent material of the hot-end components of aero-engines. Taking the semi-circular artificial crack specimen made of the powder superalloy FGH96 material as an example, each step and related technical features in the above remaining life prediction method will be further described in detail as follows:

[0078] The material parameters of the powder superalloy FGH96 material are shown in Tables 1 and 2 below:

[0079] Table 1

[0080]

[0081] Table 2

[0082]

[0083] The fatigue test data of the powder superalloy FGH96 material under a stress ratio of 0.05 are shown in Table 3 below:

[0084] Table 3

[0085]

[0086]

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

[0088] Specifically, it includes:

[0089] Obtain the material parameters of the materials constituting the hot-end components of the aero-engine; the material parameters here are the parameters in Table 1 above; based on the finite element model, set the singularity of the mesh near the crack front and apply different external loads; input the material parameters into the finite element model to obtain the J-integral at the crack front, and perform principal stress extraction to obtain the stress triaxiality at the crack front.

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

[0091] In this figure, the abscissa represents the angle between the nodes at the crack front and the crack width and the ordinate represents the J-integral. There are four curves, and each curve represents a set of typical combinations of crack length (a) and stress level (b). Specifically, it includes: solid blue line: a = 1200 μm, b = 900 MPa; dashed blue line: a = 1200 μm, b = 1200 MPa; solid red line: a = 1800 μm, b = 900 MPa; dashed red line: a = 1800 μm, b = 1000 MPa; in addition, the corresponding maximum J-integral (max) is marked on each curve, as well as the J-integral value when the angle between the nodes at the crack front and the crack width is the largest;

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

[0093] From this figure, when the crack length is the same, a higher stress level leads to a higher J-integral value. When the stress level is the same, a longer crack length also results in a higher J-integral value. And as the angle increases, the J-integral value first gradually rises to the maximum and then levels off or slightly decreases. This indicates that there is an optimal angle at which the J-integral reaches the maximum.

[0094] In addition, in finite element analysis, it is also necessary to extract the principal stress values of multiple nodes at the crack front. Based on 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 reflecting the material failure tendency under a multiaxial stress state.

[0095] Stress triaxiality calculation formula:

[0096]

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

[0098] In this embodiment S2, extrapolation processing is performed on the stress triaxiality of the nodes at the crack front to obtain the extrapolation processing result of the stress triaxiality of the nodes at the crack front;

[0099] As Figure 3 shown, the stress triaxiality of the nodes at the crack front is obtained from the stress triaxiality distribution diagrams at different ratios; where r represents the distance from the node to the crack front, σ Y represents the yield stress, and J represents the J-integral.

[0100] In this figure, the abscissa represents the ratio, r represents the distance from the node to the crack front, σ Y represents the yield stress, and J represents the J-integral; the ordinate represents the stress triaxiality The black dots in the figure represent the data points obtained by actual measurement or calculation, while the red slanted line is the fitting curve obtained by extrapolating these data points, clarifying the trend that the stress triaxiality h gradually decreases as increases.

[0101] And when tends to zero, the initial value of the stress triaxiality is approximately 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. At positions far from the crack tip, the stress state of the material becomes more uniform or closer to the uniaxial tension state.

[0102] When performing the extrapolation of stress triaxiality, first obtain the actual stress triaxiality of several nodes at the crack front using finite element analysis. Subsequently, select several typical nodes near the crack front and, based on the trend of the stress triaxiality changing with distance, construct a straight line through these known points. This straight line can assist in predicting the stress triaxiality of nodes beyond the known range. That is, use the linear relationship between the existing data points to estimate the stress triaxiality at unknown positions and achieve effective extrapolation of the stress triaxiality of external nodes.

[0103] In this embodiment S3, based on the J integral of the nodes at the crack front and the results of the stress triaxiality extrapolation, determine the representative position of the crack front in combination with a preset threshold;

[0104] Specifically, in the finite element analysis, the J integral of the nodes at the crack front and the results of the stress triaxiality extrapolation have been calculated. To determine the representative position of the crack front, first find the maximum value of the J integral among the nodes and the maximum value of the stress triaxiality extrapolation results. Then, screen out 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. These nodes can better reflect the crack propagation driving force at the crack front and the local stress state at the crack front.

[0105] Furthermore, cross-screen the above-screened nodes, select the nodes that simultaneously meet the representative conditions of the J integral and the stress triaxiality extrapolation results as the final representative positions, and select the node at the deepest part of the crack as the representative position. Finally, perform data verification on the selected representative positions to ensure the rationality and consistency of their J integral and stress triaxiality extrapolation results, and verify the rationality of the selection of the representative positions by comparing with the data of other nodes.

[0106] In this embodiment S4, generate a traditional failure assessment diagram through the J integral of the 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, which is the ordinate of the failure assessment curve; L r is the ligament yield parameter, which is the abscissa of the failure assessment curve; J e represents the elastic J integral, J ecdenotes the elastoplastic J integral, σ ref denotes the reference stress, σ 0.2 denotes the material yield strength, σ app denotes the remotely applied stress, σ0 denotes the reference yield stress, denotes the critical ligament yield parameter, σ u denotes the material tensile strength;

[0110] and when L r = 1, σ0 = σ app , satisfying:

[0111]

[0112] wherein, E denotes Young's modulus, ε ref denotes the reference strain, J e (σ app ), J ec(σapp) respectively denote the elastic J integral and the elastoplastic J integral under the remotely applied stress σ app .

[0113] In this embodiment S5, using the corresponding coordinates on the said traditional failure assessment diagram, a correlation function between the crack driving force and the ligament yield parameter under different crack sizes is constructed;

[0114] Specifically, it includes:

[0115] Based on the correlation between the J integral and the crack size a and the crack driving force f(L r ), the crack driving force f(L r ) is expressed by the crack size a:

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

[0117] wherein, f(L r (a)) denotes the crack driving force related to the crack size, and P, M, N, H denote fitting parameters;

[0118] Based on the correlation between the reference yield stress σ0 and the crack size a and the 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 the component has no crack, σ0(a = 0) = σ 0.2 , satisfying:

[0122] σ0(a) = σ 0.2 + M′a + N′a 2

[0123] where 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′, N′ represent 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 incorporating the crack size effect is generated;

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

[0126] Critical ligament yield parameter is expressed as:

[0127]

[0128] where the critical ligament yield parameter represents the magnification of the critical ligament yield parameter by times.

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

[0130] The damage tolerance failure assessment diagram includes: a meaningless region, a crack non - propagation region, a safe region, and a dangerous region.

[0131] Specifically, the meaningless region is region; the crack non - propagation region is region, where K th is the fatigue crack growth threshold; the safe region is and enclosed region; the dangerous region is or f(L r (a)) > K r region.

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

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

[0134] The calculation of the assessment curve is expressed as:

[0135]

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

[0137] Furthermore, based on the prediction result of the allowable crack size, combined with the crack propagation model, the remaining fatigue life is predicted to obtain the fatigue life prediction result;

[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 projected area of the crack in the loading direction, and c, n represent material constants.

[0141] As Figure 5 shown, based on the comparison result of the allowable crack size of the powder superalloy FGH96 material predicted by the remaining life prediction method of this embodiment with the actual size, its prediction accuracy is within the ±1.2 times dispersion band, and the error is within ±20%.

[0142] As Figure 6 shown, based on the comparison result of the remaining fatigue life of the powder superalloy FGH96 material predicted by the remaining life prediction method of this embodiment with the actual remaining life, the prediction accuracy under 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 for the hot-end components of aero-engines, such as aero-engine turbine disks. By comprehensively applying finite element analysis, failure assessment diagrams, and crack propagation models, the crack size and remaining fatigue life are accurately predicted. It fully considers the crack size effect, load conditions, and material properties, significantly improves the prediction accuracy, provides strong technical support for the damage tolerance design and in-service maintenance strategy of aero-engine hot-end components, can effectively reduce the risk of structural failure, and enhances the reliability and safety of the equipment.

[0144] The various embodiments in this specification are described in a progressive manner. Each embodiment focuses on the differences from other embodiments. For the same or similar parts among the embodiments, reference can be made to each other. For the systems disclosed in the embodiments, since they correspond to the methods disclosed in the embodiments, the description is relatively simple. For related parts, reference can be made to the description in the method section.

[0145] The above description of the disclosed embodiments enables those skilled in the art to implement or use the present invention. Various modifications to these embodiments will be obvious to those skilled in the art. The general principles defined herein can 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 these embodiments shown herein, but rather to the broadest scope consistent with the principles and novel features disclosed herein.

Claims

1. A remaining life prediction method applied to hot end components of an aeroengine, characterized in that It includes the following steps: S1. Perform finite element analysis on the hot end components of the aeroengine with cracks to determine the J integral and stress triaxiality of the nodes at the crack front under different loads; S2. Extrapolate the stress triaxiality of the nodes at the crack front to obtain the extrapolation result of the stress triaxiality of the nodes at the crack front; S3. Based on the J integral of the nodes at the crack front and the extrapolation result of the stress triaxiality, determine the representative position of the crack front in combination with a preset threshold; S4. Generate a traditional failure assessment diagram through the J integral of the representative position of the crack front; 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 yield strength of the material, σ app represents the applied stress at the far end, 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 , satisfying: where E represents Young's modulus, and ε ref represents the reference strain, and J e (σ app ) and respectively represent the elastic J-integral and elastoplastic J-integral under the distal applied stress σ app ; S5. Use the corresponding coordinates on the traditional failure assessment diagram to construct the correlation function between the crack driving force and the ligament yield parameter under different crack sizes; S6. Generate a failure assessment diagram considering crack size effect through the correlation function between the crack driving force and the ligament yield parameter under different crack sizes and loads; S7. Generate a damage tolerance failure assessment diagram based on the failure assessment diagram considering crack size effect; S8. Based on the damage tolerance failure assessment diagram, predict the allowable crack size to obtain the prediction result of the allowable crack size; S9. Based on the prediction result of the allowable crack size, combine with the crack propagation model to predict the remaining fatigue life to obtain the prediction result of the fatigue life.

2. The remaining life prediction method for hot end components of an aeroengine according to claim 1, characterized in that The S1 includes: Obtain the material parameters of the materials constituting the hot end components of the aeroengine; Based on the finite element model, set the singularity of the mesh near the crack front and apply different external loads; Input the material parameters into the finite element model to obtain the J integral of the crack front and extract the principal stress to obtain the stress triaxiality of the crack front; among them, the J integral is obtained through the J integral distribution diagram at different angles between the nodes at the crack front and the crack width, including elastic J integral and elastoplastic J integral.

3. A remaining life prediction method for hot end components of an aeroengine according to claim 1, characterized in that In S2, the stress triaxiality of the nodes at the crack front is obtained from the stress triaxiality distribution diagrams at different ratios; where r represents the distance from the node to the crack front, and σ Y represents the yield stress, and J represents the J-integral.

4. A remaining life prediction method for hot end components of an aeroengine according to claim 1, characterized in that In the S3, the representative position of the crack front includes the 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.

5. The remaining life prediction method for hot end components of an aero-engine according to claim 1, characterized in that The S5 includes: Based on the J integral and crack size a, crack driving force f(L r ) is related to the crack driving force f(L r ) to express: f(L r (a))=P+Ma+Na 2 + Ha 3 where f(L r (a)) represents the crack driving force related to the crack size, and P, M, N, and H represent fitting parameters; Based on the correlation between the reference yield stress σ0, the crack size a, and the J integral, express the ligament yield parameter Lr in terms of 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 , satisfying: 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′, N′ represent the fitting parameters.

6. 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 considering the crack size effect in S6, the ordinate of the failure assessment curve is the crack driving force f(L r (a)) related to the crack size, and the abscissa of the failure curve is the ligament yield parameter L r (a); Critical ligament yield parameter Expressed as: Among them, the critical ligament yield parameter represents the critical ligament yield parameter is magnified times.

7. A remaining life prediction method for hot end components of an aeroengine according to claim 1, characterized in that In the S7, the damage tolerance failure assessment diagram includes: a meaningless area, a non-crack propagation area, a safe area, and a dangerous area.

8. A remaining life prediction method for hot end components of an aeroengine according to claim 1, characterized in that, In the S8, the allowable crack size is determined by the intersection coordinates of the damage tolerance failure assessment curve and the assessment curve; The calculation of the assessment curve is expressed as: Among them, K r represents the ordinate of the evaluation curve, K(a) represents the stress intensity factor for different crack sizes, and K mat represents the fracture toughness, and L r (a) represents the abscissa of the evaluation curve.

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 the 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 projected area of the crack in the loading direction, and c and n represent material constants.

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