Aero-engine blade fatigue failure analysis method
By combining a real blade database and a finite element model with an air gun test system to simulate external damage, the problem of existing simulated blade designs relying on subjective judgment is solved, and efficient and low-cost high-cycle fatigue limit testing of blades is achieved.
Patent Information
- Application Number
- CN202411816597.7
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- Filing Date
- 2024-12-11
- Publication Date
- 2025-12-05
- Estimated Expiration
- 2044-12-11
AI Technical Summary
Existing simulation methods for blades lack effective external damage simulation schemes, which limits the high-cycle fatigue performance of blades. Furthermore, the design relies on subjective judgment, resulting in high costs and uncontrollable parameters.
A database was established using real blade external damage morphology data. Simulation and testing were conducted using a finite element model. External damage tests were carried out in conjunction with an air gun test system, and high-cycle fatigue strength tests were performed to ensure consistency between the simulation results and the real data.
It achieves a more realistic simulation of external damage, reduces costs, improves parameter controllability and experimental interpretability, and obtains a more accurate high-cycle fatigue limit for blades.
Smart Images
Figure CN119808467B_ABST
Abstract
Description
TECHNICAL FIELD
[0001] The present application relates to the technical field of aero-engines, in particular to an aero-engine blade fatigue failure analysis method. BACKGROUND
[0002] When the aircraft engine is in service near the ground, the strong suction of the high-speed inlet airflow often sucks foreign hard objects such as sand, rivets, stones, metal blocks and the like into the engine, which impact the fan or compressor blades of the engine, causing impact damage, and further damaging the airfoil and stress distribution of the blade, affecting the structural strength of the blade. In the domestic and foreign engine design standards, the impact damage caused by the above-mentioned hard object impact on the engine flow passage components is called foreign object damage (FOD). Since the aero-engine blade is in a high-speed rotating state during work, the relative speed between the small hard objects sucked into the engine flow passage and the blade can reach more than 350 m / s, and the hard object impact on the blade can cause notch or pit type damage at the leading edge of the blade. FOD can cause stress concentration at the damage location, and under the superimposed action of the centrifugal force load brought by high speed and the high vibration stress load caused by short-time resonance in the working process of the blade, the FOD position is prone to initiate fatigue cracks, which is an important factor limiting the high cycle fatigue (HCF) performance of the blade.
[0003] Since the real blade is expensive and has a long production cycle, the design of the simulation blade for test research is very important. In the prior art, flat plate simulation blades and dog bone simulation blades were used in the early stage, and in recent years, new types of simulation blades such as leading edge simulation blades and curved leading edge simulation blades have been developed, but the simulation blades still lack a scheme for simulating foreign object damage. SUMMARY
[0004] The purpose of the present application is to provide an aero-engine blade fatigue failure analysis method based on real foreign object damage characteristics.
[0005] Technical scheme: An aero-engine blade fatigue failure analysis method, comprising the following steps:
[0006] (1) Collecting foreign object damage topography data of real blades to form a real blade database;
[0007] (2) Using the topography and stress distribution data of the real blade as the design parameters of the simulation blade;
[0008] (3) According to the design parameters of the simulation blade, a simulation blade finite element model is established, modal calculation is performed, the natural frequencies of each order mode of the simulation blade finite element model are obtained, whether the natural frequencies of each order mode meet the set requirements is verified, if the set requirements are met, the next step is entered, if the set requirements are not met, step (2) is executed again;
[0009] (4) simulating the foreign object damage on the simulation blade finite element model to obtain a simulation result, and verifying the simulation result with the real blade database for consistency;
[0010] (5) manufacturing a real simulation blade model according to the simulation blade design parameters, and performing a foreign object damage test on the real simulation blade model according to the simulation result to obtain foreign object damage size data;
[0011] (6) performing a high-cycle fatigue test on the real simulation blade model after the foreign object damage test to obtain a high-cycle fatigue strength of the real simulation blade model.
[0012] Specifically, the foreign object damage morphology data of the real blade includes morphology data of mechanical damage and corrosion damage, the mechanical damage includes notches, tears, scratches, bending deformation, and pits, and the corrosion damage includes pitting, oxidation, and sulfuration.
[0013] Specifically, step (2) includes: performing a steady-state stress and vibration stress simulation test on the real blade to obtain stress distribution of the real blade, extracting geometric structure parameters and stress gradients of feature parts of the blade, optimizing the size of the simulation blade based on the stress gradients, and obtaining simulation blade design parameters consistent with the real blade parameters.
[0014] Specifically, in step (3), the first-order modal natural frequency of the simulation blade finite element model is set to be greater than the difference between the other order modal natural frequencies by a set value, and the maximum stress position is located at the leading edge.
[0015] Specifically, step (4) includes: performing meshing on the simulation blade finite element model, giving material parameters and element types, limiting all degrees of freedom, and impacting the three cross-section tensile pieces of the simulation blade finite element model with steel balls, concrete, and sand balls of different sizes in a dynamics analysis software to complete the foreign object damage simulation, obtain a simulation result, and verify the simulation result with the real blade database for consistency.
[0016] Specifically, in step (5), the leading edge of the real simulation blade model is impacted by an air cannon test system, and the damage part after the foreign object impact is measured to obtain foreign object damage size data.
[0017] Specifically, in step (6), a high-cycle fatigue test is performed on the real simulation blade model after the foreign object damage test by using a step method, and the high-cycle fatigue test cycle is 3x10 7 times.
[0018] Specifically, the simulation result is the size and morphology characteristics of the damage notch corresponding to the foreign object type, impact speed, and impact angle.
[0019] Specifically, the material parameters of the simulation blade finite element model are set as the material parameters of TC6 special alloy.
[0020] Specifically, the air cannon test system is an NH-10 air cannon test system.
[0021] Beneficial effects: Compared with the prior art, the significant effects of the present application are: the present method uses a simulation blade finite element model to perform impact dynamics simulation of foreign object damage, obtains the relationship between the macroscopic features of the damage notch and the foreign object type, impact speed and impact angle, and performs consistency verification with the foreign object damage morphology data of the real blade, so as to design foreign object damage tests based on real blade data rather than subjective judgment; then, the foreign object damage test conditions of the simulation blade real model are determined based on the test parameters of the simulation blade finite element model, and the simulation blade real model after damage is subjected to high-cycle fatigue test, so as to obtain the high-cycle fatigue strength of the blade, realize high-cycle fatigue limit test of the foreign object damage simulation blade which is closer to the real working condition, and the cost is lower, the parameters are highly controllable, and the method has good interpretability. BRIEF DESCRIPTION OF DRAWINGS
[0022] Figure 1 is a flow chart of the method of the present application. DETAILED DESCRIPTION
[0023] The present application will be further described below with reference to the accompanying drawings.
[0024] Please refer to Figure 1 The present application provides an aero-engine blade fatigue failure analysis method, as shown in the accompanying drawings, comprising the following steps:
[0025] Step 1: Collecting foreign object damage morphology data of real blades, foreign object damage including mechanical damage and corrosion damage, mechanical damage including notches, tears, scratches, bending deformation, pits, corrosion damage including pitting, oxidation, sulfuration, and forming a real blade database with the above foreign object damage morphology data for subsequent consistency comparison.
[0026] Step 2: Taking the morphology and stress distribution data of the real blade as the simulation blade design parameters. First, perform steady-state stress and vibration stress simulation test on the real blade to obtain the stress distribution of the blade, and extract the geometric structure parameters and stress gradient of the characteristic parts of the blade. Based on the analysis of the maximum stress position of the blade, the design of the leading edge of the simulation blade needs to be focused on, and the design suitable for bending fatigue test is obtained. Based on the stress gradient, the size of the simulation blade is optimized to obtain the simulation blade design parameters consistent with the real blade parameters.
[0027] Step 3: According to the simulation blade design parameters, a simulation blade finite element model is established in a finite element analysis software such as workbench, and modal calculation is performed to obtain the natural frequency of each order mode of the simulation blade finite element model. According to the calculation results of the natural frequency of each order mode, it is analyzed whether there is an inherent frequency of other order modes near the first order mode inherent frequency. Specifically, the difference between the first order mode inherent frequency and the inherent frequency of other order modes of the simulation blade finite element model is calculated. If it is greater than a set value, it meets the set requirement, and the maximum stress position should be located at the leading edge. At this time, it is considered that the design of the simulation blade finite element model is feasible, and the next step is entered. If it does not meet the set requirement, the simulation blade design parameters need to be adjusted again.
[0028] Step 4: The simulation blade finite element model is meshed, the material parameters and element types are given, the constraints are applied on the bottom clamp to limit all degrees of freedom, and the steel balls, concrete and sand balls of different sizes are used to impact the tension pieces of three sections of the simulation blade finite element model in a dynamics analysis software such as LS-DYNA. In some specific embodiments, the material parameters of TC6 special alloy are used as the material parameters of the simulation blade finite element model. After the external object damage simulation is completed, the simulation results are obtained, and the simulation results are consistent with the real blade database. If the score of the consistency verification is greater than a set value, it is proved that the simulation results of the external object damage simulation match the external object damage morphology data of the real blade.
[0029] Step 5: According to the simulation blade design parameters, a simulation blade real model is made, and an external object damage test (FOD) is performed on the simulation blade real model according to the simulation results. The leading edge of the simulation blade real model is impacted by an air cannon test system, and the damage position after the external object impact is measured to obtain the external object damage size data.
[0030] In some specific embodiments, the NH-10 air cannon test system is used for gas cannon high-speed impact test, and the three-dimensional body microscope is used for macroscopic measurement of the external object damage size. In one specific example, the test parameters are as follows: the material of the simulation blade real model is titanium alloy, the included angle between the center line of the leading edge of the simulation blade real model and the width direction line of the blade root is set to 20°, and to avoid the influence of the thickness change of the leading edge of the transition section on the impact damage result, the midpoint of the leading edge arc 5 mm above the maximum stress point is selected as the accurate impact position of the external object damage test.
[0031] Step 6: The simulation blade real model after the external object damage test is subjected to high cycle fatigue test (HCF). Specifically, the step-by-step method is used to perform high cycle fatigue test on the simulation blade real model after the external object damage test, and the high cycle fatigue test cycle is 3x10 7Secondly, the high cycle fatigue strength of the damaged real model of the simulation blade is obtained.
[0032] In this example, the DONGFENG digital electric vibration test system is used to perform the dwell test at the first modal natural frequency of the test piece. The strain gauge BE120-3AA produced by AVIC Electrometric is used, and the real-time data of the strain gauge during the experiment is recorded by the DH5981 dynamic strain acquisition system.
[0033] The specific test process is as follows: the strain gauges are pasted at the characteristic points of the real model of the simulation blade, the wiring is welded by electric welding, the real model of the simulation blade is installed on the vibration table and tightened. The sweep frequency test is performed to determine the resonance frequency of the real model of the simulation blade, then the dwell test is performed at the resonance frequency, and the fatigue limit of the blade is obtained by the step method.
[0034] When performing the sweep frequency test, the sweep frequency range is determined according to the modal analysis results, the sweep frequency load spectrum is written, the load is determined as 1g acceleration, and the sweep rate is 200Hz / min. The curve after the sweep frequency test is subjected to fast Fourier transform (FFT) to obtain the natural frequency. The dwell load spectrum is written at the natural frequency of the test piece to perform the dwell test, the strain data is monitored during the test, and the acceleration is adjusted to keep the measured strain at the target value. The step method is used for the dwell test, and the fatigue life is determined as N1=3×10 7 When the stress value decreases significantly in a short time, it indicates that a crack has been initiated, the cycle number N2 at this time is recorded, and the simulation blade real model is detected by the flaw detection agent. If no crack is detected, continue to apply the vibration stress and continue the step method test; if a crack is detected, analyze the corresponding test data, find the time when the strain decreases significantly, and calculate the fatigue limit of the simulation blade real model.
Claims
1. A method of fatigue failure analysis of an aeroengine blade, characterized by, The method comprises the following steps: (1) collecting the foreign object damage morphology data of real blades to form a real blade database; (2) taking the morphology and stress distribution data of the real blades as the design parameters of the simulation blades; (3) establishing a simulation blade finite element model according to the design parameters of the simulation blades, and performing modal calculation to obtain the natural frequencies of each order mode of the simulation blade finite element model, and verifying whether the natural frequencies of each order mode meet the set requirements, if yes, entering the next step, if not, performing step (2) again; (4) performing foreign object damage simulation on the simulation blade finite element model to obtain simulation results, and verifying the consistency of the simulation results with the real blade database; the simulation results are the size and morphology characteristics of the damage notch corresponding to the foreign object type, impact speed and impact angle; (5) manufacturing a real simulation blade model according to the design parameters of the simulation blades, and performing foreign object damage test on the real simulation blade model according to the simulation results to obtain foreign object damage size data; (6) performing high-cycle fatigue test on the real simulation blade model after the foreign object damage test to obtain the high-cycle fatigue strength of the real simulation blade model.
2. The aeroengine blade fatigue failure analysis method of claim 1, wherein: The foreign object damage morphology data of the real blades includes the morphology data of mechanical damage and corrosion damage, the mechanical damage includes notch, tearing, scratch, bending deformation and pit, and the corrosion damage includes pitting, oxidation and sulfuration.
3. The aeroengine blade fatigue failure analysis method of claim 1, wherein: The step (2) comprises: performing steady-state stress and vibration stress simulation test on the real blades to obtain the stress distribution of the real blades, extracting the geometric structure parameters and stress gradient of the characteristic parts of the real blades, optimizing the size of the simulation blades based on the stress gradient, and obtaining the simulation blade design parameters consistent with the real blade parameters.
4. The aeroengine blade fatigue failure analysis method of claim 1, wherein: In the step (3), the set requirements are that the difference between the first order modal natural frequency and the natural frequencies of other orders of the simulation blade finite element model is greater than a set value, and the maximum stress position is located at the leading edge.
5. The aeroengine blade fatigue failure analysis method of claim 1, wherein: The step (4) comprises: performing mesh division on the simulation blade finite element model, giving material parameters and element types, limiting all degrees of freedom, impacting the tension pieces of three sections of the simulation blade finite element model with steel balls, concrete and sand balls of different sizes in a dynamics analysis software to complete the foreign object damage simulation, obtaining the simulation results, and verifying the consistency of the simulation results with the real blade database.
6. The aeroengine blade fatigue failure analysis method of claim 1, wherein: In the step (5), the leading edge of the real simulation blade model is impacted by the foreign object by using an air cannon test system, and the damage part after the foreign object impact is measured to obtain the foreign object damage size data.
7. The aeroengine blade fatigue failure analysis method of claim 1, wherein: In the step (6), the high-cycle fatigue test is performed on the real model of the simulated blade after the foreign object damage test by using the step-by-step method, and the high-cycle fatigue test cycle is 3x10 7 times.
8. The aeroengine blade fatigue failure analysis method of claim 1, wherein: The material parameters of the simulation blade finite element model are set as the material parameters of TC6 special alloy.
9. The aeroengine blade fatigue failure analysis method of claim 6, wherein: The air cannon test system is an NH-10 air cannon test system.