Fatigue limit prediction method for surface-strengthened aero-engine blade

Abstract

The invention belongs to the technical field of aero-engine blades, and particularly discloses a surface-strengthened aero-engine blade fatigue limit prediction method, which comprises the following steps of: constructing a simulated blade finite element model of a titanium alloy simulated blade to perform first-order modal simulation analysis by analyzing a surface state evolution test result of a titanium alloy material sample after surface strengthening; and analyzing a first-order bending vibration fatigue test result of the titanium alloy simulation blade to obtain a corresponding structural stress parameter for prediction so as to obtain a fatigue limit prediction value of the surface-strengthened aero-engine blade. According to the method, the fatigue performance of the surface-strengthened blade can be accurately evaluated by adopting a blade high-cycle fatigue strength evaluation mode considering the surface state and the evolution law of the surface-strengthened blade of the aero-engine, and the problem of predicting the high-cycle fatigue strength of the blade after the surface of a complex structure is strengthened is solved; the method has important scientific significance and engineering value for manufacturing and application of aero-engine blades.

Description

Technical Field

[0001] The invention belongs to the technical field of aeroengine blades, and in particular relates to a fatigue limit prediction method for a surface-reinforced aeroengine blade. Background Art

[0002] Fan and compressor blades are the core components of aircraft engines and are mostly made of titanium alloy. Titanium alloy blades will produce high-frequency vibrations under the action of high-speed airflow and are very prone to fatigue failure. The high-cycle fatigue failure of blades is an important factor restricting the service of aircraft engines. Surface strengthening technologies such as shot peening and laser shock can introduce residual compressive stress and gradient microstructure organization on the blade surface to improve the fatigue performance of the blade, and are therefore widely used in blade manufacturing. However, due to the irregular geometric shape of the blade, the surface state such as residual stress and microstructure after surface strengthening is complex, and the surface state parameters will also evolve under the action of high-cycle fatigue loads, which brings challenges to the evaluation of the fatigue performance of surface-strengthened blades.

[0003] The current internationally used blade fatigue strength assessment model is mainly aimed at material-level tests such as plate-like parts. However, given the complex structure of aircraft engine blades, how to accurately assess the fatigue performance of surface-reinforced blades is still a very challenging problem. Summary of the invention

[0004] The purpose of the present invention is to provide a method for predicting the fatigue limit of a surface-reinforced aero-engine blade, so as to solve the above-mentioned problems existing in the prior art.

[0005] In order to achieve the above object, the present invention adopts the following technical solutions:

[0006] In a first aspect, a method for predicting fatigue limit of a surface-strengthened aero-engine blade is provided, comprising:

[0007] Obtain the test results of the surface state evolution of the fatigue specimens of the surface-strengthened titanium alloy material, and obtain the fatigue strength test results of the titanium alloy material before and after surface strengthening;

[0008] Determine the fatigue danger point of the titanium alloy simulated blade surface after strengthening based on the surface state evolution test results, and determine the strengthening factor based on the fatigue strength test results of the titanium alloy material;

[0009] Constructing a simulated blade finite element model of a titanium alloy simulated blade, wherein the simulated blade finite element model includes a plurality of mesh units, and performing a first-order modal simulation using the simulated blade finite element model to determine a simulated dangerous point and a far-field point, as well as a stress ratio between the simulated dangerous point and the far-field point;

[0010] Obtain the first-order bending vibration fatigue test results of the titanium alloy simulated blade, and determine the test stress corresponding to the simulated danger point when the titanium alloy simulated blade reaches the fatigue limit according to the first-order bending vibration fatigue test results;

[0011] The unstrengthened stress corresponding to the far-field point is calculated using the test stress corresponding to the simulated dangerous point and the stress ratio between the simulated dangerous point and the far-field point, and the strengthened stress corresponding to the far-field point is calculated using the unstrengthened stress corresponding to the far-field point and the strengthening factor;

[0012] Extract the simulation stress parameters of the first-order modal simulation of the blade finite element model, and calculate the average stress corresponding to the fatigue danger point using the critical distance method based on the simulation stress parameters;

[0013] The fatigue limit of the titanium alloy blade after surface strengthening is estimated based on the strengthening stress corresponding to the far-field point, the stress ratio between the simulation danger point and the far-field point, and the average stress corresponding to the fatigue danger point.

[0014] In a possible design, the fatigue danger point of the titanium alloy simulated blade after surface strengthening is determined according to the surface state evolution test results, including:

[0015] According to the surface state evolution test results, the calibrated stress σ' and interpolated residual stress σ' corresponding to each point of the fatigue specimen of the surface-strengthened titanium alloy are determined. res and interpolated fatigue limit σ' -1 ;

[0016] According to the calibrated stress σ' and interpolated residual stress σ' corresponding to each point res and interpolated fatigue limit σ' -1 Calculate the evolution law parameter corresponding to each point, the evolution law parameter = σ' -1 / (σ'+σ' res );

[0017] Constructing a two-dimensional scatter plot using the evolution law parameters corresponding to each point, displaying the two-dimensional scatter plot, and selecting a point where the evolution law parameters in the two-dimensional scatter plot have an inflection point in response to a first operation instruction of the user;

[0018] The selected points are used as fatigue danger points after the titanium alloy simulated blade surface is strengthened.

[0019] In a possible design, determining the strengthening factor according to the fatigue strength test results of the titanium alloy material includes:

[0020] The fatigue limit of the titanium alloy material before and after strengthening is determined under a set stress ratio according to the fatigue strength test results of the titanium alloy material. The strengthening factor is obtained by dividing the fatigue limit of the titanium alloy material after strengthening by the fatigue limit before strengthening.

[0021] In a possible design, the first-order modal simulation is performed using a simulated blade finite element model to determine a simulated dangerous point and a far-field point, as well as a stress ratio between the simulated dangerous point and the far-field point, including:

[0022] Use the simulated blade finite element model to perform first-order modal simulation and display the simulation result model;

[0023] In response to the second operation instruction of the user, a far-field point and a simulated dangerous point are selected in the simulation result model, and a simulated stress σ of the far-field point in the simulation result model is determined. A_sim , and the simulated stress σ of the simulated dangerous point in the simulation result model B_sim ;

[0024] Using the simulated stress σ at the far-field point A_sim and the simulated stress σ at the simulated dangerous point B_sim Calculate the stress ratio γ between the simulated dangerous point and the far-field point, where γ = σ B_sim / σ A_sim .

[0025] In a possible design, the calculation of the unstrengthened stress corresponding to the far-field point using the test stress corresponding to the simulated dangerous point and the stress ratio between the simulated dangerous point and the far-field point includes:

[0026] The unstrengthened stress corresponding to the far-field point is obtained by dividing the test stress corresponding to the simulated dangerous point by the stress ratio between the simulated dangerous point and the far-field point.

[0027] In a possible design, the calculating of the strengthened stress corresponding to the far-field point using the unstrengthened stress corresponding to the far-field point and the strengthening factor includes:

[0028] The unenhanced stress corresponding to the far-field point is multiplied by the enhancement factor to obtain the enhanced stress corresponding to the far-field point.

[0029] In a possible design, extracting the simulated stress parameters of the first-order modal simulation of the blade finite element model, and calculating the average stress corresponding to the fatigue danger point by using the critical distance method based on the simulated stress parameters, includes:

[0030] Determination of residual stress σ after surface strengthening of titanium alloy simulated blade res , determine the critical distance r of all center points from the fatigue danger point c The grid cells in the range are located at the critical distance r from the fatigue hazard point. c The grid cells within the range are used as target cells, where:

[0031] r c =1.54l 0

[0032]

[0033] M is the set geometric correction coefficient, σ -1 The material fatigue strength of the simulated blade of a given titanium alloy, ΔK th is the set fatigue threshold stress intensity factor;

[0034] Determine the absolute value of the maximum principal stress of each target unit, and determine the average stress σ of all target units based on the absolute value of the maximum principal stress of each target unit C , and using the strengthening factor and the mean stress σ C and residual stress σ res Calculate the superimposed stress σ of each target unit i ,in:

[0035] σ i =ασ C +σ res

[0036] α is the reinforcement factor, i represents the serial number of the target unit;

[0037] Using the superimposed stress σ of each target unit i Calculate the average stress σ corresponding to the fatigue danger point aver ,in:

[0038]

[0039] η is the set bending effect correction coefficient, N is the number of target units, V i is the unit volume of target unit i.

[0040] In a possible design, the fatigue limit of the titanium alloy blade after surface strengthening is estimated based on the strengthening stress corresponding to the far-field point, the stress ratio between the simulated danger point and the far-field point, and the average stress corresponding to the fatigue danger point, including:

[0041] Substitute the strengthening stress corresponding to the far-field point, the stress ratio between the simulation danger point and the far-field point, and the average stress corresponding to the fatigue danger point into the preset surface strengthening blade fatigue limit prediction formula for calculation to obtain the fatigue limit of the aircraft engine blade surface strengthened titanium alloy material. The surface strengthening blade fatigue limit prediction formula is:

[0042]

[0043] Among them, σ A'_sim It represents the strengthening stress corresponding to the far-field point, and γ represents the stress ratio between the simulation danger point and the far-field point.

[0044] In a second aspect, a fatigue limit prediction system for a surface-strengthened aero-engine blade is provided, comprising a first acquisition unit, a result determination unit, a modal simulation unit, a second acquisition unit, a first calculation unit, a second calculation unit, and a limit estimation unit, wherein:

[0045] The first acquisition unit is used to acquire the surface state evolution test results of the fatigue specimen of the surface-strengthened titanium alloy material, and to acquire the fatigue strength test results of the titanium alloy material before and after the surface strengthening;

[0046] A result determination unit is used to determine the fatigue danger point of the titanium alloy simulated blade surface after strengthening according to the surface state evolution test results, and to determine the strengthening factor according to the fatigue strength test results of the titanium alloy material;

[0047] A modal simulation unit is used to construct a simulated blade finite element model of a titanium alloy simulated blade, wherein the simulated blade finite element model includes a plurality of mesh units, and the simulated blade finite element model is used to perform a first-order modal simulation to determine a simulated dangerous point and a far-field point, as well as a stress ratio between the simulated dangerous point and the far-field point;

[0048] The second acquisition unit is used to obtain the first-order bending vibration fatigue test results of the titanium alloy simulated blade, and determine the test stress corresponding to the simulated dangerous point when the titanium alloy simulated blade reaches the fatigue limit according to the first-order bending vibration fatigue test results;

[0049] A first calculation unit is used to calculate the unstrengthened stress corresponding to the far-field point by using the test stress corresponding to the simulated dangerous point and the stress ratio between the simulated dangerous point and the far-field point, and calculate the strengthened stress corresponding to the far-field point by using the unstrengthened stress corresponding to the far-field point and the strengthening factor;

[0050] The second calculation unit is used to extract the simulation stress parameters of the first-order modal simulation of the blade finite element model, and calculate the average stress corresponding to the fatigue danger point by using the critical distance method based on the simulation stress parameters;

[0051] The limit estimation unit is used to estimate the fatigue limit of the titanium alloy blade after surface strengthening based on the strengthening stress corresponding to the far-field point, the stress ratio between the simulation danger point and the far-field point, and the average stress corresponding to the fatigue danger point.

[0052] In a third aspect, a fatigue limit prediction system for a surface-strengthened aero-engine blade is provided, comprising:

[0053] A memory for storing instructions;

[0054] A processor is used to read the instructions stored in the memory and execute any one of the methods described in the first aspect according to the instructions.

[0055] In a fourth aspect, a computer-readable storage medium is provided, wherein instructions are stored on the computer-readable storage medium, and when the instructions are executed on a computer, the computer executes any one of the methods described in the first aspect. In addition, a computer program product is provided, and when the computer program product is executed on a computer, the computer executes any one of the methods described in the first aspect.

[0056] Beneficial effect: The present invention adopts a blade high-cycle fatigue strength assessment method that takes into account the surface state and evolution law of aircraft engine blades after surface reinforcement. It can accurately assess the fatigue performance of surface-reinforced blades and solve the problem of predicting the high-cycle fatigue strength of blades with complex structures after surface reinforcement. It has important scientific significance and engineering value for the manufacture and application of aircraft engine blades. BRIEF DESCRIPTION OF THE DRAWINGS

[0057] 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 some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0058] Figure 1 This is a schematic diagram of the steps of the method in Example 1 of the present invention;

[0059] Figure 2 It is the two-dimensional scatter plot in Example 1 of the present invention;

[0060] Figure 3 It is a schematic diagram of the simulation result model in Example 1 of the present invention;

[0061] Figure 4 Schematic diagram of the system structure in Example 3 of the present invention. DETAILED DESCRIPTION

[0062] It should be noted that the description of these embodiments is used to help understand the present invention, but does not constitute a limitation of the present invention. The specific structures and functional details disclosed herein are only used to describe the exemplary embodiments of the present invention. However, the present invention can be embodied in many alternative forms, and it should not be understood that the present invention is limited to the embodiments set forth herein.

[0063] It should be understood that unless otherwise clearly specified and limited, the corresponding terms should be understood in a broad sense. For example, "connection" can be a fixed connection, a detachable connection, or an integral connection; it can be a direct connection, or an indirect connection through an intermediate medium, or it can be the internal communication of two elements. For ordinary technicians in this field, the specific meanings of the above terms in the embodiments can be understood according to specific circumstances.

[0064] In the following description, certain details are provided to facilitate a complete understanding of the example embodiments. However, it will be appreciated by those of ordinary skill in the art that the example embodiments may be implemented without these certain details. For example, devices may be shown in block diagrams to avoid obscuring the examples with unnecessary details. In other embodiments, well-known processes, structures, and techniques may not be shown in unnecessary detail to avoid obscuring the embodiments.

[0065] Embodiment 1:

[0066] This embodiment provides a method for predicting fatigue limit of surface-strengthened aero-engine blades, which can be applied to corresponding evaluation systems, such as Figure 1 As shown, the method comprises the following steps:

[0067] S1. Obtain the test results of the surface state evolution of the fatigue specimen of the surface-strengthened titanium alloy material, and obtain the fatigue strength test results of the titanium alloy material before and after surface strengthening.

[0068] In specific implementation, the surface state evolution test of the fatigue specimen of the surface-strengthened titanium alloy material (the same titanium alloy material as the aero-engine blade, such as TC17 titanium alloy material) can be carried out to obtain the corresponding surface state evolution test results. The fatigue strength test of the titanium alloy material surface of the aero-engine blade before and after laser shock strengthening can be carried out to obtain the corresponding fatigue strength test results.

[0069] S2. Determine the fatigue danger point of the titanium alloy simulated blade surface after strengthening based on the surface state evolution test results, and determine the strengthening factor based on the fatigue strength test results of the titanium alloy material.

[0070] In specific implementation, the surface state evolution test results contain the evolution law of surface states such as residual stress and microhardness under fatigue load, based on which the fatigue danger point of the titanium alloy simulated blade after surface strengthening can be determined. For example, the surface state evolution test results of the surface strengthened TC17 titanium alloy sample are shown in Table 1 after interpolation processing:

[0071] Table 1

[0072]

[0073]

[0074] Based on the surface state evolution test results, the calibrated stress σ' and interpolated residual stress σ' corresponding to each point can be determined. res and interpolated fatigue limit σ' -1 ; Then according to the calibration stress σ' and interpolation residual stress σ' corresponding to each point res and interpolated fatigue limit σ'-1 Calculate the evolution law parameter corresponding to each point, the evolution law parameter = σ' -1 / (σ'+σ' res );Then use the evolution law parameters corresponding to each point to construct Figure 2 The two-dimensional scatter plot shown in the figure displays the two-dimensional scatter plot. The user can see from the two-dimensional scatter plot that the value of the evolution law parameter first decreases rapidly with the increase of depth, and then decreases slowly after reaching the inflection point. Although there is no obvious minimum point in the middle, since cracks always initiate from close to the surface, fatigue cracks are more likely to initiate at the inflection point of the image than at deeper interiors. This point can be considered as the dangerous point of the surface-strengthened TC17 titanium alloy simulated blade. At this time, the point depth is 80mm. The user can calibrate the point through the first operation instruction. In response to the user's first operation instruction, the system selects the point where the evolution law parameter in the two-dimensional scatter plot has an inflection point, and uses the selected point as the fatigue dangerous point after the titanium alloy simulated blade surface is strengthened.

[0075] The fatigue limit of the titanium alloy material before and after laser shock strengthening under the set stress ratio conditions can be determined according to the fatigue strength test results of the titanium alloy material. The fatigue limit of the titanium alloy material after laser shock strengthening is divided by the fatigue limit before laser shock strengthening to obtain the strengthening factor α. For example, when the stress ratio is 0.1, the fatigue limits of the TC17 titanium alloy material before and after laser shock strengthening are 510MPa and 565.71MPa, respectively. After conversion by the Goodman formula, when the stress ratio is -1, the fatigue limits of the TC17 titanium alloy material before and after strengthening are 707.73MPa and 802.65MPa, respectively. The fatigue limit after strengthening of 802.65MPa at this time is divided by the fatigue limit before strengthening of 707.73MPa, and the strengthening factor α can be calculated to be 1.13.

[0076] S3. Construct a simulated blade finite element model of a titanium alloy simulated blade, wherein the simulated blade finite element model includes a plurality of mesh units, and use the simulated blade finite element model to perform a first-order modal simulation to determine a simulated danger point and a far-field point, as well as a stress ratio between the simulated danger point and the far-field point.

[0077] In specific implementation, a simulated blade finite element model of a titanium alloy simulated blade can be constructed according to the design size of the titanium alloy simulated blade, and the simulated blade finite element model includes a number of mesh units. Then, the simulated blade finite element model is used to perform a first-order modal simulation, and the simulation result model is displayed. For example, the simulation result model is as follows: Figure 3 As shown, the user can determine the far-field point A and the simulated dangerous point B according to the displayed simulation result model. In response to the user's second operation instruction, the system selects the far-field point A and the simulated dangerous point B in the simulation result model, and determines the simulated stress σ of the far-field point A in the simulation result model.A_sim , and the simulated stress σ of the simulated dangerous point B in the simulation result model B_sim ; Finally, the simulated stress σ at the far-field point A is used A_sim and the simulated stress σ of the simulated dangerous point B B_sim Calculate the stress ratio γ between the simulated dangerous point and the far-field point, where γ = σ B_sim / σ A_s i m .

[0078] S4. Obtain the first-order bending vibration fatigue test results of the titanium alloy simulated blade, and determine the test stress corresponding to the simulated danger point when the titanium alloy simulated blade reaches the fatigue limit based on the first-order bending vibration fatigue test results.

[0079] During specific implementation, a first-order bending vibration fatigue test can be carried out on the titanium alloy simulated blade to obtain the first-order bending vibration fatigue test results of the unreinforced blade. The system determines the test stress corresponding to the simulated danger point when the titanium alloy simulated blade reaches the fatigue limit based on the first-order bending vibration fatigue test results.

[0080] S5. Calculate the unstrengthened stress corresponding to the far-field point using the test stress corresponding to the simulated dangerous point and the stress ratio between the simulated dangerous point and the far-field point, and calculate the strengthened stress corresponding to the far-field point using the unstrengthened stress corresponding to the far-field point and the strengthening factor.

[0081] In specific implementation, the system divides the test stress corresponding to the simulated dangerous point by the stress ratio γ between the simulated dangerous point and the far-field point to obtain the unenhanced stress corresponding to the far-field point, and then multiplies the unenhanced stress corresponding to the far-field point by the enhancement factor α to obtain the enhanced stress σ corresponding to the far-field point. A'_sim .

[0082] S6. Extract the simulation stress parameters of the first-order modal simulation of the blade finite element model, and calculate the average stress corresponding to the fatigue danger point using the critical distance method based on the simulation stress parameters.

[0083] In the specific implementation, after carrying out the numerical simulation of the first-order vibration mode of the simulated blade, all the principal stress data and Mises stress data of each grid unit are output and imported into MATLAB. The simulated stress parameters of the first-order modal simulation of the finite element model of the simulated blade are extracted through MATLAB, including the absolute value of the maximum principal stress and the residual stress, etc., and the critical distance method is used to calculate the average stress corresponding to the fatigue danger point based on the simulated stress parameters. The whole process includes:

[0084] Determination of residual stress σ after surface strengthening of titanium alloy simulated blade res , determine the critical distance r of all center points from the fatigue danger point cThe grid cells in the range are located at the critical distance r from the fatigue hazard point. c The grid cells within the range are used as target cells, where:

[0085] r c =1.54l 0

[0086]

[0087] M is the set geometric correction coefficient, which is generally 1, σ -1 The material fatigue strength of the simulated blade of a given titanium alloy, ΔK th For the fatigue threshold stress intensity factor set, for TC17 titanium alloy, it can be set as

[0088] Determine the absolute value of the maximum principal stress of each target unit, and determine the average stress σ of all target units based on the absolute value of the maximum principal stress of each target unit C For example, the mean absolute value of the maximum principal stress of each target unit is calculated based on the absolute value of the maximum principal stress of the target unit, and the mean absolute value of the maximum principal stress is multiplied by the proportional coefficient β between the given actual stress measurement value of the enhanced blade and the numerical simulation stress of the first-order vibration mode (such as the simulation stress σ of the far-field point A). A_sim , whose actual stress measurement value is σ A'_test , then β=σ A'_test / σ A_sim ), and the average stress σ of all target units is obtained C Then, using the strengthening factor and the mean stress σ C and residual stress σ res Calculate the superimposed stress σ of each target unit i ,in:

[0089] σ i =ασ C +σ res

[0090] α is the reinforcement factor, i represents the serial number of the target unit;

[0091] Using the superimposed stress σ of each target unit i Calculate the average stress σ corresponding to the fatigue danger point aver ,in:

[0092]

[0093] η is the set bending effect correction coefficient, N is the number of target units, V i is the unit volume of target unit i.

[0094] S7. Estimate the fatigue limit of the titanium alloy blade after surface strengthening based on the strengthening stress corresponding to the far-field point, the stress ratio between the simulation danger point and the far-field point, and the average stress corresponding to the fatigue danger point.

[0095] In specific implementation, the system substitutes the strengthening stress corresponding to the far-field point, the stress ratio between the simulation danger point and the far-field point, and the average stress corresponding to the fatigue danger point into the preset surface strengthening blade fatigue limit prediction formula for calculation, and obtains the fatigue limit of the aircraft engine titanium alloy blade after surface strengthening. The surface strengthening blade fatigue limit prediction formula is:

[0096]

[0097] Among them, σ A'_sim It represents the strengthening stress corresponding to the far-field point, and γ represents the stress ratio between the simulation danger point and the far-field point.

[0098] In order to verify the prediction accuracy of the method of this embodiment, bending vibration fatigue tests were carried out on 6 laser shock strengthened TC17 titanium alloy simulated blades to measure their actual fatigue limits. The measurement results are shown in Table 2 below:

[0099] Table 2

[0100]

[0101] The results show that the fatigue limit test value σ of 6 laser shock strengthened titanium alloy blades e Compared with the predicted values, the prediction errors of the six laser shock strengthened titanium alloy blades were all within 10%, reflecting good prediction accuracy.

[0102] Embodiment 2:

[0103] This embodiment provides a fatigue limit prediction system for a surface-strengthened aircraft engine blade, comprising a first acquisition unit, a result determination unit, a modal simulation unit, a second acquisition unit, a first calculation unit, a second calculation unit, and a limit estimation unit, wherein:

[0104] The first acquisition unit is used to acquire the surface state evolution test results of the fatigue specimen of the surface-strengthened titanium alloy material, and to acquire the fatigue strength test results of the titanium alloy material before and after the surface strengthening;

[0105] A result determination unit is used to determine the fatigue danger point of the titanium alloy simulated blade surface after strengthening according to the surface state evolution test results, and to determine the strengthening factor according to the fatigue strength test results of the titanium alloy material;

[0106] A modal simulation unit is used to construct a simulated blade finite element model of a titanium alloy simulated blade, wherein the simulated blade finite element model includes a plurality of mesh units, and the simulated blade finite element model is used to perform a first-order modal simulation to determine a simulated dangerous point and a far-field point, as well as a stress ratio between the simulated dangerous point and the far-field point;

[0107] The second acquisition unit is used to obtain the first-order bending vibration fatigue test results of the titanium alloy simulated blade, and determine the test stress corresponding to the simulated dangerous point when the titanium alloy simulated blade reaches the fatigue limit according to the first-order bending vibration fatigue test results;

[0108] A first calculation unit is used to calculate the unstrengthened stress corresponding to the far-field point by using the test stress corresponding to the simulated dangerous point and the stress ratio between the simulated dangerous point and the far-field point, and calculate the strengthened stress corresponding to the far-field point by using the unstrengthened stress corresponding to the far-field point and the strengthening factor;

[0109] The second calculation unit is used to extract the simulation stress parameters of the first-order modal simulation of the blade finite element model, and calculate the average stress corresponding to the fatigue danger point by using the critical distance method based on the simulation stress parameters;

[0110] The limit estimation unit is used to estimate the fatigue limit of the titanium alloy blade after surface strengthening based on the strengthening stress corresponding to the far-field point, the stress ratio between the simulation danger point and the far-field point, and the average stress corresponding to the fatigue danger point.

[0111] Embodiment 3:

[0112] This embodiment provides a fatigue limit prediction system for surface-strengthened aero-engine blades. Figure 4 As shown, at the hardware level, it includes:

[0113] Data interface, used to establish data connection between the processor and the external data terminal;

[0114] A memory for storing instructions;

[0115] The processor is used to read the instructions stored in the memory and execute the surface-reinforced aircraft engine blade fatigue limit prediction method in Example 1 according to the instructions.

[0116] Optionally, the system further includes an internal bus, through which the processor, the memory and the data interface can be interconnected, and the internal bus can be an ISA (Industry Standard Architecture) bus, a PCI (Peripheral Component Interconnect) bus or an EISA (Extended Industry Standard Architecture) bus, etc. The bus can be divided into an address bus, a data bus, a control bus, etc.

[0117] The memory may include, but is not limited to, random access memory (RAM), read only memory (ROM), flash memory, first input first output (FIFO) and / or first in last out (FILO) memory, etc. The processor may be a general-purpose processor, including a central processing unit (CPU), a network processor (NP), etc.; it may also be a digital signal processor (DSP), an application specific integrated circuit (ASIC), a field programmable gate array (FPGA) or other programmable logic devices, discrete gate or transistor logic devices, discrete hardware components.

[0118] Embodiment 4:

[0119] This embodiment provides a computer-readable storage medium, on which instructions are stored, and when the instructions are executed on a computer, the computer executes the surface-strengthened aircraft engine blade fatigue limit prediction method in Embodiment 1. The computer-readable storage medium refers to a carrier for storing data, which may include but is not limited to a floppy disk, an optical disk, a hard disk, a flash memory, a USB flash drive, and / or a memory stick, etc., and the computer may be a general-purpose computer, a special-purpose computer, a computer network, or other programmable devices.

[0120] This embodiment also provides a computer program product, which, when running on a computer, executes the fatigue limit prediction method of the surface-strengthened aerospace engine blade in Embodiment 1. The computer may be a general-purpose computer, a special-purpose computer, a computer network or other programmable device.

[0121] Finally, it should be noted that the above description is only a preferred embodiment of the present invention and is not intended to limit the protection scope of the present invention. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

Claims

1. A method for predicting fatigue limit of surface-strengthened aero-engine blades, characterized in that: include: Obtain the test results of the surface state evolution of the fatigue specimens of the surface-strengthened titanium alloy material, and obtain the fatigue strength test results of the titanium alloy material before and after surface strengthening; Determine the fatigue danger point of the titanium alloy simulated blade surface after strengthening based on the surface state evolution test results, and determine the strengthening factor based on the fatigue strength test results of the titanium alloy material; Constructing a simulated blade finite element model of a titanium alloy simulated blade, wherein the simulated blade finite element model includes a plurality of mesh units, and performing a first-order modal simulation using the simulated blade finite element model to determine a simulated dangerous point and a far-field point, as well as a stress ratio between the simulated dangerous point and the far-field point; Obtain the first-order bending vibration fatigue test results of the titanium alloy simulated blade, and determine the test stress corresponding to the simulated danger point when the titanium alloy simulated blade reaches the fatigue limit according to the first-order bending vibration fatigue test results; The unstrengthened stress corresponding to the far-field point is calculated using the test stress corresponding to the simulated dangerous point and the stress ratio between the simulated dangerous point and the far-field point, and the strengthened stress corresponding to the far-field point is calculated using the unstrengthened stress corresponding to the far-field point and the strengthening factor; Extract the simulation stress parameters of the first-order modal simulation of the blade finite element model, and calculate the average stress corresponding to the fatigue danger point using the critical distance method based on the simulation stress parameters; The fatigue limit of the titanium alloy blade after surface strengthening is estimated based on the strengthening stress corresponding to the far-field point, the stress ratio between the simulation danger point and the far-field point, and the average stress corresponding to the fatigue danger point.

2. A method for predicting fatigue limit of surface-strengthened aero-engine blades according to claim 1, characterized in that: The method of determining the fatigue danger point of the titanium alloy simulated blade surface after strengthening according to the surface state evolution test results includes: According to the surface state evolution test results, the calibrated stress σ' and interpolated residual stress σ' corresponding to each point of the fatigue specimen of the surface-strengthened titanium alloy are determined. res and interpolated fatigue limit σ' -1 ; According to the calibrated stress σ' and interpolated residual stress σ' corresponding to each point res and interpolated fatigue limit σ' -1 Calculate the evolution law parameter corresponding to each point, the evolution law parameter = σ' -1 / (σ'+σ' res ); Constructing a two-dimensional scatter plot using the evolution law parameters corresponding to each point, displaying the two-dimensional scatter plot, and selecting a point where the evolution law parameters in the two-dimensional scatter plot have an inflection point in response to a first operation instruction of the user; The selected points are used as fatigue danger points after the titanium alloy simulated blade surface is strengthened.

3. The method for predicting fatigue limit of surface-strengthened aero-engine blades according to claim 1, characterized in that: Determining the strengthening factor according to the fatigue strength test results of the titanium alloy material includes: The fatigue limit of the titanium alloy material before and after strengthening is determined under a set stress ratio according to the fatigue strength test results of the titanium alloy material. The strengthening factor is obtained by dividing the fatigue limit of the titanium alloy material after strengthening by the fatigue limit before strengthening.

4. The method for predicting fatigue limit of surface-strengthened aero-engine blades according to claim 1, characterized in that: The method of using the simulated blade finite element model to perform first-order modal simulation to determine the simulated dangerous point and the far-field point, as well as the stress ratio between the simulated dangerous point and the far-field point, includes: Use the simulated blade finite element model to perform first-order modal simulation and display the simulation result model; In response to the second operation instruction of the user, a far-field point and a simulated dangerous point are selected in the simulation result model, and a simulated stress σ of the far-field point in the simulation result model is determined. A_sim , and the simulated stress σ of the simulated dangerous point in the simulation result model B_sim ; Using the simulated stress σ at the far-field point A_sim and the simulated stress σ at the simulated dangerous point B_sim Calculate the stress ratio γ between the simulated dangerous point and the far-field point, where γ = σ B_sim / σ A_sim .

5. The method for predicting fatigue limit of surface-strengthened aero-engine blades according to claim 1, characterized in that: The method of calculating the unstrengthened stress corresponding to the far-field point by using the test stress corresponding to the simulated dangerous point and the stress ratio between the simulated dangerous point and the far-field point includes: The unstrengthened stress corresponding to the far-field point is obtained by dividing the test stress corresponding to the simulated dangerous point by the stress ratio between the simulated dangerous point and the far-field point.

6. The method for predicting fatigue limit of surface-strengthened aero-engine blades according to claim 1, characterized in that: The method of calculating the strengthened stress corresponding to the far-field point by using the unstrengthened stress corresponding to the far-field point and the strengthening factor comprises: The unenhanced stress corresponding to the far-field point is multiplied by the enhancement factor to obtain the enhanced stress corresponding to the far-field point.

7. A method for predicting fatigue limit of surface-strengthened aero-engine blades according to claim 1, characterized in that: The extracting of the simulated stress parameters of the first-order modal simulation of the blade finite element model and calculating the average stress corresponding to the fatigue danger point by using the critical distance method based on the simulated stress parameters include: Determination of residual stress σ after surface strengthening of titanium alloy simulated blade res , determine the critical distance r of all center points from the fatigue danger point c The grid cells in the range are located at the critical distance r from the fatigue hazard point. c The grid cells within the range are used as target cells, where: r c ＝1.54l0 M is the set geometric correction coefficient, σ -1 The material fatigue strength of the simulated blade of a given titanium alloy, ΔK th is the set fatigue threshold stress intensity factor; Determine the absolute value of the maximum principal stress of each target unit, and determine the average stress σ of all target units based on the absolute value of the maximum principal stress of each target unit C , and using the strengthening factor and the mean stress σ C and residual stress σ res Calculate the superimposed stress σ of each target unit i ,in: s i =as C +s res α is the reinforcement factor, i represents the serial number of the target unit; Using the superimposed stress σ of each target unit i Calculate the average stress σ corresponding to the fatigue danger point aver ,in: η is the set bending effect correction coefficient, N is the number of target units, V i is the unit volume of target unit i.

8. A method for predicting fatigue limit of surface-strengthened aero-engine blades according to claim 7, characterized in that: The method of estimating the fatigue limit of the titanium alloy blade after surface strengthening according to the strengthening stress corresponding to the far-field point, the stress ratio between the simulation danger point and the far-field point, and the average stress corresponding to the fatigue danger point includes: Substitute the strengthening stress corresponding to the far-field point, the stress ratio between the simulation danger point and the far-field point, and the average stress corresponding to the fatigue danger point into the preset surface strengthening blade fatigue limit prediction formula for calculation to obtain the fatigue limit of the aircraft engine blade surface strengthened titanium alloy material. The surface strengthening blade fatigue limit prediction formula is: Among them, σ A'_sim It represents the strengthening stress corresponding to the far-field point, and γ represents the stress ratio between the simulation danger point and the far-field point.

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