CN113945388B - Truncated test method for vibration fatigue test of aero-engine blade - Google Patents
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- G01M15/00—Testing of engines
- G01M15/14—Testing gas-turbine engines or jet-propulsion engines
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- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
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Abstract
The invention belongs to the technical field of aeroengine strength tests, and discloses a truncated test method for an aeroengine blade vibration fatigue test, which comprises the following steps: s1, acquiring a mounting mode of a whole machine blade, carrying out finite element analysis on vibration characteristics of the whole machine blade, and calculating to obtain a main stress direction of the blade in the whole machine state; s2, based on a test state blade mounting mode, carrying out finite element analysis on vibration characteristics of the test state blade, and calculating to obtain a main stress direction of the test state blade; s3, comparing, analyzing and judging the difference of the main stress directions of the blade in the whole machine state and the test state, shortening the length direction of the blade, and recalculating the main stress directions of the blade until the main stress directions are equal to each other; s4, carrying out truncated processing on the blade, and carrying out vibration fatigue test on the blade based on the truncated processing. The invention not only can keep the stress direction of the test state and the real working state consistent, but also can shorten the test period and the cost.
Description
Technical Field
The invention belongs to the technical field of aero-engine strength tests, and particularly relates to an engine blade vibration fatigue test method, in particular to a truncated test method for an aero-engine blade vibration fatigue test.
Background
The aeroengine is known as a pearl on a modern industrial crown, is a heart of an airplane, is a high-speed rotating thermodynamic mechanical device, is one of the most common faults of the engine based on the fact that the aeroengine is in a high-temperature, high-pressure and high-speed severe environment, and the falling block, the crack and the like caused by the blade vibration fault threaten the safety of the whole engine, so that a large number of blade vibration fatigue tests can be carried out in the development process of the engine, the median fatigue strength of the blade with the confidence of 95% and the survival rate of 50% can be determined, the structural design, material selection and manufacturing process rationality of the blade can be evaluated, and technical guarantees are provided for the safe and reliable operation of the blade in the whole life cycle of the engine.
At present, the vibration fatigue test of the blade mainly adopts a constraint mode that one end is clamped and fixed and the other end is free, and the method is as follows: yang Weixin, li Yan, wang Ping A novel aero-engine blade fatigue test method [ J ]. Noise and vibration control, 2017,37 (5): 214-218 and literature: li Sailu, cheng Li, liu Jingyuan the plate blade nonlinear vibration and fatigue test [ J ]. The university of air force engineering university (natural science edition), 2017,18 (5): 1-6, are different from the real installation mode of the blade in the engine, especially for the stator blade (the real installation mode is two-end constraint and the test state is one-end constraint), the maximum point of the test vibration fatigue stress and the difference between the stress direction and the real engine state are caused by the large difference of the two, and the vibration fatigue test cannot better reflect the actual vibration mode and vibration mechanism of the real blade.
Meanwhile, in the blade vibration fatigue test, the test cycle times are higher, and are generally more than 10 7(3*107 times, and some of the test cycle times even reach 10 9 times, but in order to improve the thrust-weight ratio, the current advanced aero-engine is pursued to lighten the structure, the blade is generally made of materials with lower density, the first-order natural frequency of the blade is relatively low (the first-order frequency is less than 100 Hz), the time consumed for carrying out one blade vibration fatigue test is very long, the time is less than tens of hours, the time is more than days, and the one vibration fatigue needs to finish the blade test with the specified number, so the time spent for one blade vibration fatigue test is less than one month, the time spent for more than months, and the consumed labor cost and the test cost are both relatively high.
In order to better solve the difference of maximum stress points and stress directions between the checking position and the real using state checking position in the vibration fatigue test of the blade, and in order to shorten the test period, the existing blade fatigue test method needs to be improved. The invention provides a technical invention of a vibration fatigue shortening scheme of an engine blade.
Disclosure of Invention
The invention overcomes the defects existing in the prior art, and solves the technical problems that: the method for shortening the vibration fatigue test of the aero-engine blade improves the test efficiency and shortens the test time on the premise of improving the test accuracy.
In order to solve the technical problems, the invention adopts the following technical scheme: a truncated test method for an aeroengine blade vibration fatigue test comprises the following steps:
s1, acquiring a mounting mode of a whole machine blade, setting boundary conditions consistent with the operation of the whole machine, carrying out finite element analysis on vibration characteristics of the blade in the whole machine state, acquiring vibration stress parameters of the blade in the whole machine state, and calculating according to the vibration stress parameters to obtain a main stress direction of the blade in the whole machine state;
S2, based on a test state blade mounting mode, carrying out finite element analysis on vibration characteristics of the test state blade to obtain vibration stress parameters of the test state blade, and calculating according to the vibration stress parameters to obtain a main stress direction of the test state blade;
s3, comparing and analyzing to judge whether the main stress directions of the blades are equal in the whole machine state and the test state, if not, truncating the blades in the length direction, returning to the step S2, recalculating the main stress directions of the blades, and if so, entering the step S4;
S4, cutting the blade according to the length of the blade when the main directions of the stresses are equal, and performing vibration fatigue test on the blade after cutting.
In the step S1 and the step S2, the obtained vibration stress parameters include the natural vibration frequency, the vibration mode, the relative vibration stress and the stress direction of the blade.
In the step S1 and the step S2, when the main stress direction of the blade is calculated, the first-order vibration mode of the blade is taken as a study object.
In the step S1 and the step S2, the method for calculating the main stress direction of the blade is as follows:
and drawing a Morse circle according to the stress state of the blade and a Morse circle calculation formula, and determining a stress main direction according to the Morse circle.
The calculation formula for determining the main stress direction according to the Morse circle is as follows:
wherein, Representing the angle between the principal stress and the x-axis, τ xy represents the shear stress, σ x represents the stress along the x-direction, and σ y represents the stress along the y-direction.
In the step S3, the shortening amount of the next blade is adjusted according to the difference between the main stress direction in the test state of the last blade and the main stress direction in the complete machine state.
In the step S3, the specific method for shortening the blade in the length direction includes:
firstly setting an initial chopping quantity to truncate the blade, calculating the main stress direction after truncating, comparing the main stress direction with the state of the whole machine, then adjusting the chopping quantity for a plurality of times, repeatedly calculating the main stress direction after truncating to obtain the relation between the chopping quantity and the main stress direction, and adjusting the chopping quantity according to the relation until the chopping quantity of the blade consistent with the main stress direction in the state of the whole machine is obtained.
Compared with the prior art, the invention has the following beneficial effects:
The invention provides a truncated test method for an aeroengine blade vibration fatigue test, which overcomes the defects that in the prior art, the main stress direction of the maximum point of vibration stress of a stator blade is different from the real state of the engine to a certain extent, the test state cannot better reflect the technical defect of the complete machine state, the main stress direction of the stress is kept consistent through analysis, the truncated blade is subjected to the fatigue test, the main stress direction of the maximum vibration point is consistent with the complete machine state, the complete machine state can be effectively simulated during the test, the blade vibration frequency can be improved, the test period is shortened, and the test cost is effectively saved.
Drawings
FIG. 1 is a technical flow chart of a truncated test method for an aero-engine blade vibration fatigue test provided by an embodiment of the invention;
FIG. 2 is an initial three-dimensional model of a blade with an upper shaft secured to a casing by bolts and a lower shaft inserted into an inner ring during engine assembly;
FIG. 3 is a schematic diagram of a Moire circle from which the principal direction of stress can be derived directly for a given stress state;
FIG. 4 is a schematic view of a spatial state (three-way stress) Moire circle defining three principal stresses σ 1>σ2>σ3;
FIG. 5 is a schematic illustration of a blade truncation scheme achieved according to principal direction of stress consistency in an embodiment of the present invention;
FIG. 6 is a schematic view of a blade selected during a fatigue test in which the upper end of the shaft is fixedly constrained and the lower end is free;
FIG. 7 is a graphical representation of the relative vibratory stress distribution of an adjustable blade obtained by finite element analysis;
FIG. 8 is a diagram showing stress direction of the whole machine state obtained by finite element analysis;
FIG. 9 is a schematic diagram of the stress direction (one-end anchoring) of the experimental state obtained by finite element analysis;
FIG. 10 is a schematic view of the direction of vibratory stress after blade shortening.
Detailed Description
For the purpose of making the objects, technical solutions and advantages of the embodiments of the present invention more clear, the technical solutions in the embodiments of the present invention will be clearly and completely described below, and it is apparent that the described embodiments are some embodiments of the present invention, but not all embodiments; all other embodiments, which can be made by those skilled in the art based on the embodiments of the invention without making any inventive effort, are intended to be within the scope of the invention.
The method is mainly realized by means of a blade vibration theory, a Moire circle theory in material mechanics and large commercial finite element software, an engine blade shortening scheme is determined by a finite element simulation technology, and the content comprises analysis of an engine complete machine state blade mounting mode and vibration characteristic analysis, so that data such as a first order vibration stress maximum point position, a stress main direction and the like of the blade are obtained; carrying out vibration characteristic analysis of the blade in a fatigue test state to obtain a maximum point of vibration stress, a direction of the maximum stress, a vibration mode and the like; comparing the stress maximum point position, the corresponding stress main direction, the vibration frequency and other data under the two states, and finding out the difference of the two; and carrying out vibration characteristic analysis of a blade shortening scheme by adopting a heuristic method, researching the relation between the shortening position and the stress direction of the maximum point, finding out the optimal shortening position, taking the optimal shortening position as the shortening position of the final test blade, and carrying out vibration fatigue test by adopting the shortened blade.
First, the theory of blade vibration and the cause of resonance are analyzed.
(1) Blade vibration theory.
The rotor and stator blade of the aeroengine is easy to generate resonance in operation, and according to a mechanical vibration theory, the undamped free vibration has the following motion equation:
Mü+I-P＝0；(1)
wherein: m is mass, u is acceleration, I is internal force, and P is external force applied.
In modal analysis, the internal forces include both motion (e.g., damping, etc.) and structural deformation. The natural frequency can be determined by considering the dynamic response of the unloaded structure (p=0), the equation of motion of which is:
Mü+I＝0；(2)
For undamped systems, i=ku, then there is:
Mü+Ku＝0；(3)
Wherein: k is a stiffness matrix and u is displacement.
The solution of the equation is of the form:
u＝φeiwt；(4)
Wherein: phi and w are hypothetical variables.
And (5) carrying out a motion equation to obtain a characteristic value:
Kφ＝λMφ；(5)
let λ=ω 2, equation (5) can be converted into:
det(K-λM)＝0；(6)
let lambda i be one of the eigenvalues, equation (6) can be written as:
(K-λi M)φi＝0,i＝1,2,…n；(7)
Wherein: n is the number of degrees of freedom in the finite element calculation model; let lambda j be the jth eigenvalue, its square root omega j be the natural frequency of the jth mode of the structure (Natural Frequency), and phi j be the corresponding jth eigenvector (Eigenvector). The eigenvector, mode Shape, also known as Mode Shape, is the deformed Shape of the structure vibrating in the j-th order Mode.
Modal analysis is a typical solution eigenvalue problem, and the inherent frequency of multi-order vibration and the corresponding modal vector are obtained by solving eigenvalue equations.
Resonance refers to the phenomenon that when the frequency excited by a structure approaches a certain order natural frequency of the system, or the amplitude of the system increases significantly. At resonance, the energy exciting the input mechanical system is maximum, and the obvious vibration mode of the system is called displacement resonance.
(2) And (5) analyzing resonance reasons.
Under certain conditions, the natural frequency of the component is a fixed value, and the excitation frequency is closely related to external factors such as excitation factors, environment and the like; when the excitation frequency is close to the frequency margin of the natural frequency, the member resonates. The frequency margin calculation formula is shown in formula (8), and for different components, the frequency margin range delta i is different, the smaller the delta i is, the greater the possibility of resonance is, and specifically, the following steps are:
Wherein: Δ i is a frequency margin value, f Excitation device is an excitation frequency, and f Fixing device is a natural frequency.
(3) And drawing a moire circle and determining the stress direction.
According to the stress state of the blade and a Moire circle calculation formula, a Moire circle can be drawn, and taking a plane stress state as an example, the origin of the Moire circle is (a, 0), and the radius is R.
Wherein:
Wherein: a is an origin coordinate, σ x represents stress in the x direction, σ y represents stress in the y direction, and τ xy is a shear stress.
The included angle between the principal stress sigma 1 and the x-axis isThe calculation formula of the included angle is as follows:
After the stress circle is determined, the main stress direction (the included angle between the first main stress and the x direction) can be determined.
Example 1
As shown in FIG. 1, the embodiment of the invention provides a truncated test method for an aeroengine blade vibration fatigue test, which comprises the following steps:
S1, analyzing vibration characteristics of the whole machine state blade and determining stress directions: and obtaining the installation mode of the whole machine blade, setting boundary conditions consistent with the operation of the whole machine, carrying out finite element analysis on the vibration characteristics of the blade in the whole machine state by a finite element analysis method, obtaining the vibration stress parameters of the blade in the whole machine state, and calculating according to the vibration stress parameters to obtain the main stress direction of the blade in the whole machine state.
Taking an air inlet casing adjustable blade of a military engine with a small bypass ratio as an example, setting boundary conditions (full constraint at the upper end and rotatable and immovable at the lower end) consistent with the working state of the whole engine by means of large commercial finite element software and pretreatment screening software, carrying out finite element analysis on the vibration characteristics of the blade in the whole engine state, obtaining the inherent vibration frequency, vibration mode, relative vibration stress, stress direction and the like of the blade, taking a first-order vibration mode as a research object, and recording the position, stress direction, vibration mode and other results of the maximum value of the vibration stress in detail. And obtaining the main stress direction of the whole machine state according to the stress analysis of the blade and the Moire circle calculation formulas of figures 3 and 4.
For the spatial three-dimensional stress state, the stress contrast analysis can be simplified into a plane stress state, and for the plane stress state, as can be seen from fig. 3, the main stress direction can be directly obtained under the condition that the two-dimensional stress is known.
In this embodiment, the method for calculating the principal stress direction of the blade includes: and drawing a Morse circle according to the stress state of the blade and a Morse circle calculation formula, and determining a stress main direction according to the Morse circle. Specifically, when determining the principal direction of stress from the Morse circle, the calculation may be performed according to equation (10).
S2, analyzing vibration characteristics of the blade in a test state and determining stress directions: and carrying out finite element analysis on the vibration characteristics of the blade in the test state based on the installation mode of the blade in the test state by using a finite element analysis method to obtain vibration stress parameters of the blade in the test state, and calculating to obtain the main stress direction of the blade in the test state.
Based on the actual test state (one end of the blade is fixedly constrained), the vibration characteristics of the blade are analyzed, a first-order vibration mode is taken as a research object, and the results of the position of the maximum point of the vibration stress, the stress direction, the vibration mode and the like are recorded in detail, wherein the main stress direction is in the first-order mode.
Specifically, in step S2, the blade stress main direction is calculated in the same manner as in step S1.
S3, comparing and analyzing to judge whether the main stress directions of the blades are equal in the whole machine state and the test state, if not, setting a shortening amount, shortening the blades in the length direction, returning to the step S2, and recalculating the main stress directions of the blades, and if so, entering the step S4.
The first-order vibration mode, the maximum vibration stress and the stress direction of the blade in the whole machine state and the test state are compared and analyzed, the difference of the maximum stress position and the main stress direction is analyzed, the difference of the main stress directions of the two is obtained, and the fact that the test state cannot accurately simulate the vibration stress direction of the real working state of the blade is explained.
In order to obtain the blade shortening scheme, the test effectively simulates the state of the whole machine, and the modal analysis of the blade shortening scheme is developed. Because the magnitude of the truncation is unknown, heuristics are used for analysis and a stress principal direction calculation applet is developed. Firstly, carrying out certain amount of shortening on the blade, calculating the difference between the main stress direction after shortening and the state of the whole machine, then carrying out gradual adjustment according to specific conditions to obtain the relation between the main stress direction and the main stress direction, and finally obtaining the blade shortening scheme consistent with the state of the whole machine as the blade test shortening scheme.
Specifically, in step S3, in the cycle of the shortening test, the shortening amount of the next blade is adjusted according to the difference between the main stress direction in the test state of the previous blade and the main stress direction in the complete machine state.
Therefore, in this embodiment, the specific method for blade shortening is as follows: firstly setting an initial chopping quantity to truncate the blade, calculating the main stress direction after truncating, comparing the main stress direction with the state of the whole machine, then adjusting the chopping quantity for a plurality of times, repeatedly calculating the main stress direction after truncating to obtain the relation between the chopping quantity and the main stress direction, and adjusting the chopping quantity according to the relation until the chopping quantity of the blade consistent with the main stress direction in the state of the whole machine is obtained.
S4, cutting the blade according to the length of the blade when the main directions of the stresses are equal, and performing vibration fatigue test on the blade after cutting.
In this embodiment, for different blades, due to inconsistent structural forms and sizes, the truncated sizes are inconsistent and are determined according to specific conditions; because the vibration frequency of the blade is effectively improved after the blade is shortened, the cycle number of the vibration fatigue test of the blade is generally more than 10 7 times, the test period can be shortened due to the frequency improvement, and the test cost can be effectively saved.
Specifically, taking an air inlet casing adjustable blade as an example, the maximum value of the first-order relative vibration stress of the blade is located at the transition fillet position of the rotating shaft at the upper end of the blade and the blade body, as shown in fig. 7, the stress direction of the blade in the whole machine state is shown in fig. 8, the stress direction (the blade is not truncated) in the prior art scheme is shown in fig. 9, and certain difference exists between the main stress directions of the maximum point of the vibration stress in the two states, and the difference is larger as shown by calculation and analysis of Morse circles.
As shown in FIG. 10, the blade vibration mode analysis of different shortening schemes is performed by adopting a heuristic method, so that the shortening scheme with the direction consistent with the maximum stress point of the whole machine state can be found out and used as a final test scheme. Compared with the non-truncated scheme, the truncated blade has the first section vibration frequency 3 times that of the non-truncated scheme, the test time is shortened to 1/3 of the original test time, and the truncated scheme not only keeps the stress direction of the test state and the stress direction of the real working state consistent, but also shortens the test period and the cost, and is an effective scheme worthy of popularization and application.
In summary, the invention provides a truncated test method for an aeroengine blade vibration fatigue test, which overcomes the technical problems that the main stress direction of the maximum point of vibration stress of a stator blade is different from the real state of an engine to some extent, the test state cannot better reflect the state of the whole machine, and the main stress direction of the maximum vibration point of the blade in the test state is consistent with the state of the whole machine through finite element analysis, so that the test process can effectively simulate the state of the whole machine, and the provided technical scheme can improve the vibration frequency of the blade, shorten the test period and effectively save the test cost. Assuming that the first-order vibration frequency of the blade is 100Hz and the specified cycle number is 10 7, the time required for completing the test of one blade is 10 7/100/3600=27.8 hours, the time required for the test is 10 7/120/3600=23.1 hours and the time required for the test of one blade is 4.7 hours, for the test, a plurality of blades are generally required to pass the test, and the time required for the whole test is 4.7 x (wherein x is the specified number of the blades), so that the gas power cost and the labor cost of the test can be saved, the design and development process of the project can be accelerated, and the social benefit is huge.
Finally, it should be noted that: the above embodiments are only for illustrating the technical solution of the present invention, and not for limiting the same; although the invention has been described in detail with reference to the foregoing embodiments, it will be understood by those of ordinary skill in the art that: the technical scheme described in the foregoing embodiments can be modified or some or all of the technical features thereof can be replaced by equivalents; such modifications and substitutions do not depart from the spirit of the invention.
Claims (7)
1. A truncated test method for an aeroengine blade vibration fatigue test is characterized by comprising the following steps:
s1, acquiring a mounting mode of a whole machine blade, setting boundary conditions consistent with the operation of the whole machine, carrying out finite element analysis on vibration characteristics of the blade in the whole machine state, acquiring vibration stress parameters of the blade in the whole machine state, and calculating according to the vibration stress parameters to obtain a main stress direction of the blade in the whole machine state;
S2, based on a test state blade mounting mode, carrying out finite element analysis on vibration characteristics of the test state blade to obtain vibration stress parameters of the test state blade, and calculating according to the vibration stress parameters to obtain a main stress direction of the test state blade;
s3, comparing and analyzing to judge whether the main stress directions of the blades are equal in the whole machine state and the test state, if not, truncating the blades in the length direction, returning to the step S2, recalculating the main stress directions of the blades, and if so, entering the step S4;
S4, cutting the blade according to the length of the blade when the main directions of the stresses are equal, and performing vibration fatigue test on the blade after cutting.
2. The method according to claim 1, wherein the vibration stress parameters obtained in the step S1 and the step S2 include a natural vibration frequency, a vibration mode, a relative vibration stress and a stress direction of the blade.
3. The method for shortening the vibration fatigue test of the aero-engine blade according to claim 1, wherein in the step S1 and the step S2, the first-order vibration mode of the blade is taken as a study object when the stress main direction of the blade is calculated.
4. The method for shortening the vibration fatigue test of the aero-engine blade according to claim 1, wherein in the step S1 and the step S2, the method for calculating the main stress direction of the blade is as follows:
and drawing a Morse circle according to the stress state of the blade and a Morse circle calculation formula, and determining a stress main direction according to the Morse circle.
5. The truncated test method for the vibration fatigue test of the aero-engine blade according to claim 4, wherein the calculation formula for determining the principal direction of stress according to the Morse circle is as follows:
wherein, Representing the angle between the principal stress and the x-axis, τ xy represents the shear stress, σ x represents the stress along the x-direction, and σ y represents the stress along the y-direction.
6. The method according to claim 1, wherein in the step S3, the amount of shortening of the next blade is adjusted according to the difference between the main stress direction in the last blade test state and the main stress direction in the complete machine state.
7. The method for shortening the vibration fatigue test of the aero-engine blade according to claim 1, wherein in the step S3, the specific method for shortening the blade in the length direction is as follows:
firstly setting an initial chopping quantity to truncate the blade, calculating the main stress direction after truncating, comparing the main stress direction with the state of the whole machine, then adjusting the chopping quantity for a plurality of times, repeatedly calculating the main stress direction after truncating to obtain the relation between the chopping quantity and the main stress direction, and adjusting the chopping quantity according to the relation until the chopping quantity of the blade consistent with the main stress direction in the state of the whole machine is obtained.
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| CN115292924B (en) * | 2022-07-29 | 2023-06-02 | 中国航发沈阳发动机研究所 | Design method of high-cycle fatigue test run load spectrum of whole aeroengine |
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