Aero-engine blade fatigue failure analysis method

Abstract

The invention provides an aero-engine blade fatigue failure analysis method, which comprises the following steps of: firstly, collecting foreign object damage morphology data of a real blade to form a real blade database, then obtaining simulation blade design parameters according to morphology and stress distribution of the real blade, further establishing a simulation blade finite element model, and performing modal calculation to obtain a simulation blade finite element model; verifying whether the modal inherent frequency of each order meets set requirements, if the set requirements are met, performing foreign object damage analogue simulation on the simulated blade finite element model to obtain an analogue simulation result, performing consistency verification on the analogue simulation result and a real blade database, finally manufacturing a simulated blade real model, and continuously performing an FOD test and an HCF test on the simulated blade real model to obtain a simulated blade model; and obtaining the fatigue limit of the damaged simulation blade real model. According to the method, the foreign object damage morphology data of the real blade is taken as a reference, the high-cycle fatigue limit test of the foreign object damage simulation blade closer to the real working condition is realized, the cost is low, the parameter height is controllable, and the method has good interpretability.

Description

Technical Field

[0001] The invention relates to the technical field of aeroengines, and in particular to a fatigue failure analysis method for aeroengine blades. Background Art

[0002] When an aircraft engine is in service near the ground, the strong suction force generated by the high-speed intake airflow often sucks foreign hard objects such as sand, rivets, stones, and metal blocks into the engine, hitting the fan or compressor blades of the engine, causing impact damage, thereby destroying the blade shape and stress distribution, and affecting the structural strength of the blade. In domestic and international engine design standards, the impact damage caused by the above-mentioned hard objects to the engine flow path components is called foreign object damage (FOD). Since the aircraft engine blades are in a high-speed rotating state during operation, the relative speed between the small hard objects sucked into the engine flow path and the blades can reach more than 350m / s. The impact of hard objects on the blades will cause notches or pits on the leading edge of the blades. FOD will cause stress concentration at the damaged position. Under the superposition of the centrifugal force load caused by the high speed and the high vibration stress load caused by the short-term resonance of the blade during operation, fatigue cracks are very likely to initiate at the FOD position, which is also an important factor limiting the high cycle fatigue (HCF) performance of the blades.

[0003] Since real blades are expensive and have a long production cycle, the design of simulated blades used for experimental research is very important. In the existing technology, simulated blades used in the early days were flat simulated blades and dog-bone simulated blades. In recent years, new types of simulated blades such as leading edge simulated blades and curved leading edge simulated blades have been developed. However, simulated blades still lack solutions for simulating damage from foreign objects. Summary of the invention

[0004] Purpose of the invention: The purpose of the invention is to provide an aero-engine blade fatigue failure analysis method based on real foreign object damage characteristics.

[0005] Technical solution: A method for analyzing fatigue failure of an aero-engine blade, comprising the following steps:

[0006] (1) Collect the morphological data of foreign object damage on real blades and form a real blade database;

[0007] (2) Using the real blade morphology and stress distribution data as the simulated blade design parameters;

[0008] (3) According to the simulated blade design parameters, a simulated blade finite element model is established, and modal calculation is performed to obtain the natural frequencies of each order of the simulated blade finite element model, and to verify whether the natural frequencies of each order of the modals meet the set requirements. If the set requirements are met, the next step is entered; if not, step (2) is performed again;

[0009] (4) Performing foreign object damage simulation on the simulated blade finite element model to obtain simulation results, and verifying the consistency of the simulation results with the real blade database;

[0010] (5) making a simulated blade real model according to the simulated blade design parameters, and conducting a foreign object damage test on the simulated blade real model according to the simulation results to obtain foreign object damage size data;

[0011] (6) A high-cycle fatigue test is performed on the simulated blade real model after the external object damage test to obtain the high-cycle fatigue strength of the simulated blade real model.

[0012] Specifically, the morphological data of foreign object damage of real blades include morphological data of mechanical damage and corrosion damage. Mechanical damage includes notches, tears, scratches, bending deformations, and pits. Corrosion damage includes pitting, oxidation, and sulfidation.

[0013] Specifically, step (2) includes: performing steady-state stress and vibration stress simulation tests on a real blade to obtain the stress distribution of the real blade, and extracting the geometric structure parameters and stress gradients of characteristic parts of the blade, optimizing the size of the simulated blade based on the stress gradient, and obtaining simulated blade design parameters that are consistent with the real blade parameters.

[0014] Specifically, in step (3), the setting requirement is that the difference between the first-order modal natural frequency and the other-order modal natural frequencies of the simulated blade finite element model is greater than a set value, and the maximum stress position is located at the leading edge.

[0015] Specifically, step (4) includes: meshing the simulated blade finite element model, giving material parameters and unit types, restricting all degrees of freedom, using steel balls, concrete and sand balls of different sizes to impact the tensile parts of three sections of the simulated blade finite element model in the dynamic analysis software, completing the external object damage simulation, obtaining the simulation results, and verifying the consistency of the simulation results with the real blade database.

[0016] Specifically, in step (5), an air cannon test system is used to perform a foreign object impact on the leading edge of a simulated blade real model, and then the damaged 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 simulated blade real model after the foreign object damage test using the step method, and the number of high cycle fatigue test cycles is 3×10 7 Second-rate.

[0018] Specifically, the simulation results are the type of foreign object, impact velocity, impact angle and the size and morphological characteristics of the corresponding damage notch.

[0019] Specifically, the material parameters of the simulated blade finite element model are set to the material parameters of TC6 special alloy.

[0020] Specifically, the air cannon test system is an NH-10 air cannon test system.

[0021] Beneficial effect: Compared with the prior art, the significant effect of the present invention is: this method uses a simulated blade finite element model to perform impact dynamics simulation of foreign object damage, obtains the relationship between the macroscopic characteristics of the damage notch and the foreign object type, impact velocity, and impact angle, and verifies the consistency with the foreign object damage morphology data of the real blade, so that the foreign object damage test is designed based on the real blade data rather than subjective judgment; then the foreign object damage test conditions of the simulated blade real model are determined based on the test parameters of the simulated blade finite element model, and then the high-cycle fatigue test is performed on the damaged simulated blade real model, so as to obtain the high-cycle fatigue strength of the blade, and realizes the high-cycle fatigue limit test of the foreign object damage simulated blade that is closer to the real working condition, and the cost is low, the parameters are highly controllable, and it has good interpretability. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] Figure 1 It is a flow chart of the method of the present invention. DETAILED DESCRIPTION

[0023] The present invention is further described below in conjunction with the accompanying drawings.

[0024] See also Figure 1 As shown, the present invention provides a method for analyzing fatigue failure of an aero-engine blade, comprising the following steps:

[0025] Step 1: Collect the morphological data of foreign object damage on real blades. Foreign object damage includes mechanical damage and corrosion damage. Mechanical damage includes notches, tears, scratches, bending deformation, and pits. Corrosion damage includes pitting, oxidation, and sulfidation. The above morphological data of foreign object damage are combined into a real blade database for subsequent consistency comparison.

[0026] Step 2: Use the morphology and stress distribution data of the real blade as the design parameters of the simulated blade. First, conduct steady-state stress and vibration stress simulation tests 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. From the analysis of the maximum stress position of the blade, it is necessary to focus on the design of the leading edge of the simulated blade and the design suitable for bending fatigue testing. Based on the stress gradient, the size of the simulated blade is optimized to obtain the simulated blade design parameters consistent with the real blade parameters.

[0027] Step 3: According to the simulated blade design parameters, a simulated blade finite element model is established in finite element analysis software such as workbench, and modal calculation is performed to obtain the natural frequencies of each order of the simulated blade finite element model. According to the calculation results of the natural frequencies of each order of the modal, it is analyzed whether there are natural frequencies of other orders of modal near the first-order modal natural frequency. Specifically, the difference between the first-order modal natural frequency and the natural frequencies of other orders of modal of the simulated blade finite element model is calculated. If it is greater than the set value, the set requirements are met. At the same time, the maximum stress position should be located at the leading edge. At this time, the design of the simulated blade finite element model is considered feasible and enters the next step. If it does not meet the set requirements, the simulated blade design parameters need to be readjusted.

[0028] Step 4: Mesh the simulated blade finite element model, give material parameters and unit types, impose constraints on the bottom clamp, limit all degrees of freedom, use steel balls, concrete and sand balls of different sizes to impact the tensile parts of three sections of the simulated blade finite element model in dynamic analysis software such as LS-DYNA, and in some specific implementations, use the material parameters of TC6 special alloy as the material parameters of the simulated blade finite element model; after completing the foreign object damage simulation, obtain the simulation results, and verify the consistency of the simulation results with the real blade database. If the consistency verification score is greater than the set value, it proves that the simulation results obtained by the foreign object damage simulation match the foreign object damage morphology data of the real blade.

[0029] Step 5: Make a simulated blade real model according to the simulated blade design parameters, and conduct a foreign object damage test (FOD) on the simulated blade real model according to the simulation results. Use an air cannon test system to perform foreign object impact on the leading edge of the simulated blade real model, and then measure the damaged parts after the foreign object impact to obtain the foreign object damage size data.

[0030] In some specific embodiments, a NH-10 air cannon test system is used to conduct a gas cannon high-speed impact test, and then a three-dimensional stereo microscope is used to perform macroscopic measurement of the size of the foreign object damage. In a specific example, the test parameters used are: the material of the simulated blade real model is titanium alloy, the angle between the leading edge center line and the blade root width direction line of the simulated blade real model is set to 20°, and in order to avoid the influence of the leading edge thickness change of the transition section on the impact damage result, the midpoint of the leading edge arc 5mm upward from the maximum stress point is selected as the accurate impact position of the foreign object damage test.

[0031] Step 6: A high cycle fatigue test (HCF) is performed on the simulated blade real model after the foreign object damage test. Specifically, the step method is used to perform a high cycle fatigue test on the simulated blade real model after the foreign object damage test. The number of cycles of the high cycle fatigue test is 3×10 7The high cycle fatigue strength of the real model of the simulated blade after damage was obtained.

[0032] In this example, a dwell test was carried out using the Donlim digital electric vibration test system at the first-order modal natural frequency of the specimen. A strain gauge from AVIC Electro-Mechanical Test, model BE120-3AA, was used, and the real-time data of the strain gauge during the experiment was recorded using the DH5981 dynamic strain acquisition system.

[0033] The specific test process is as follows: paste the strain gauge at the characteristic point of the simulated blade real model, weld the wires with electric welding, install the simulated blade real model on the vibration table and tighten it. Perform a sweep frequency test to determine the resonance frequency of the simulated blade real model, then perform a dwell test at its resonance frequency, and obtain the fatigue limit of the blade by the step method.

[0034] When performing a frequency sweep test, the frequency sweep range is determined based on the modal analysis results, and the frequency sweep load spectrum is compiled. The load is set to 1g acceleration and the frequency sweep rate is 200Hz / min. The curve after the frequency sweep is subjected to a fast Fourier transform (FFT) to obtain the natural frequency. A dwell load spectrum is compiled at the natural frequency of the specimen to perform a dwell test. During the test, the strain data is monitored and the acceleration is adjusted to keep the measured strain at the target value. The dwell test uses the step method, and the fatigue life is set to N1=3×10 7 . When the stress value drops significantly in a short period of time, it means that cracks have been initiated. The number of cycles N2 at this time is recorded, and the real model of the simulated blade is tested with a flaw detector. If no cracks are detected, continue to apply vibration stress and continue the step method test; if cracks are detected, analyze the corresponding test data, find the time when the strain is significantly reduced, and calculate the fatigue limit of the real model of the simulated blade.

Claims

1. A method for analyzing fatigue failure of an aero-engine blade, characterized in that: The following steps are involved: (1) Collect the morphological data of foreign object damage on real blades and form a real blade database; (2) Using the real blade morphology and stress distribution data as the simulated blade design parameters; (3) According to the simulated blade design parameters, a simulated blade finite element model is established, and modal calculation is performed to obtain the natural frequencies of each order of the simulated blade finite element model, and to verify whether the natural frequencies of each order of the modals meet the set requirements. If the set requirements are met, the next step is entered; if not, step (2) is performed again; (4) Performing foreign object damage simulation on the simulated blade finite element model to obtain simulation results, and verifying the consistency of the simulation results with the real blade database; (5) making a simulated blade real model according to the simulated blade design parameters, and conducting a foreign object damage test on the simulated blade real model according to the simulation results to obtain foreign object damage size data; (6) A high-cycle fatigue test is performed on the simulated blade real model after the external object damage test to obtain the high-cycle fatigue strength of the simulated blade real model.

2. The method for analyzing fatigue failure of an aeroengine blade according to claim 1, characterized in that: 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 deformations, and pits. The corrosion damage includes pitting, oxidation, and sulfidation.

3. The method for analyzing fatigue failure of an aeroengine blade according to claim 1, characterized in that: The step (2) includes: performing steady-state stress and vibration stress simulation tests on the real blade to obtain the stress distribution of the real blade, and extracting the geometric structure parameters and stress gradients of the characteristic parts of the real blade, optimizing the size of the simulated blade based on the stress gradient, and obtaining simulated blade design parameters that are consistent with the real blade parameters.

4. The method for analyzing fatigue failure of an aeroengine blade according to claim 1, characterized in that: In the step (3), the setting requirement is that the difference between the first-order modal natural frequency and the other-order modal natural frequencies of the simulated blade finite element model is greater than a set value, and the maximum stress position is located at the leading edge.

5. The method for analyzing fatigue failure of an aeroengine blade according to claim 1, characterized in that: The step (4) includes: meshing the simulated blade finite element model, giving material parameters and unit types, limiting all degrees of freedom, using steel balls, concrete and sand balls of different sizes to impact the tensile parts of three sections of the simulated blade finite element model in the dynamic analysis software, completing the foreign object damage simulation, obtaining the simulation results, and verifying the consistency of the simulation results with the real blade database.

6. The method for analyzing fatigue failure of an aeroengine blade according to claim 1, characterized in that: In the step (5), an air cannon test system is used to perform a foreign object impact on the leading edge of a simulated blade real model, and then the damaged part after the foreign object impact is measured to obtain foreign object damage size data.

7. The method for analyzing fatigue failure of an aeroengine blade according to claim 1, characterized in that: In step (6), a high cycle fatigue test is performed on the simulated blade real model after the foreign object damage test using the step method. The number of cycles of the high cycle fatigue test is 3×10 7 Second-rate.

8. The method for analyzing fatigue failure of an aeroengine blade according to claim 1, characterized in that: The simulation results include the type of foreign object, impact velocity, impact angle and the size and morphology characteristics of the corresponding damage notch.

9. The method for analyzing fatigue failure of an aeroengine blade according to claim 1, characterized in that: The material parameters of the simulated blade finite element model are set to the material parameters of TC6 special grade alloy.

10. The method for analyzing fatigue failure of an aeroengine blade according to claim 6, characterized in that: The air cannon test system is an NH-10 air cannon test system.

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