CN118167679A - Method for diagnosing damage position of rotor of transonic compressor through aeroelastic analysis - Google Patents
Method for diagnosing damage position of rotor of transonic compressor through aeroelastic analysis Download PDFInfo
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Abstract
The invention discloses a method for diagnosing the damage position of a rotor of a transonic compressor by aeroelastic analysis, and relates to the technical field of damage inspection; the invention starts from the real response process of the compressor rotor, integrates static/dynamic pneumatic elastic solving, replaces static solving with unsteady pushing to improve the solving precision of the thermal state (under the working rotating speed) appearance, takes the thermal state appearance as a reference model of dynamic pneumatic elastic solving, further improves the solving precision of the dynamic pneumatic elasticity, carries out parameterization modeling on damaged blades on the basis, and diagnoses the damage position, the damage range and even the damage degree by analyzing the pneumatic elastic response characteristics of the damaged blades.
Description
Technical Field
The invention relates to the technical field of damage inspection, in particular to a method for diagnosing a damage position of a transonic compressor rotor through aeroelastic analysis.
Background
The structural integrity of the compressor rotor is critical to the performance and efficiency of the engine. However, rotor blades are subjected to high speed rotation and high aerodynamic loads during operation, exposing them to significant mechanical stresses and fatigue during operation, potential fatigue cracking, which in turn leads to blade breakage events, and the prior art in the field of turbine engine maintenance and rotor health monitoring relies primarily on conventional methods such as non-destructive testing (NDT) and visual inspection to assess the condition of the compressor rotor. There are some limitations to these approaches:
1. Image recognition based engine fault diagnosis, image recognition methods typically fail to detect subsurface damage or other invisible problems, such as internal damage caused by aeroelastic effects. In addition, turbine engine operating environments are often complex conditions of high temperature, high pressure, and high speed, which can cause image quality to be disturbed, thereby reducing the accuracy of the diagnosis, requiring more manual intervention to verify the results.
2. Based on non-destructive testing (NDT) of acoustic emission techniques, acoustic emission techniques cannot provide detailed information about internal problems of the material, such as subsurface damage or internal fatigue damage caused by the aeroelastic effect. In addition, the turbine engine operating environment is often very noisy, filled with high frequency noise. These environmental factors can interfere with the results of the acoustic emission test, reducing its accuracy.
In addition, there is a vibration-based damage analysis strategy, but at present, only the natural frequency of the structure itself is often detected, and the mutual coupling between pneumatic structures is ignored, so that the existing diagnosis method is insensitive to the damage position and cannot predict the potential aeroelastic accident.
In view of the foregoing, there remains a need in the art for a more accurate, real-time and non-invasive diagnostic method that can determine the location of fatigue damage in a compressor rotor.
Disclosure of Invention
Aiming at the defects of the prior art, the invention provides a method for diagnosing the damage position of a rotor of a transonic compressor by aeroelastic analysis, which solves the technical problems of limited accuracy and sensitivity of the existing damage inspection technology in the field.
In order to achieve the above purpose, the invention is realized by the following technical scheme: the method for diagnosing the damage position of the rotor of the transonic compressor by aeroelastic analysis specifically comprises the following steps:
S1, subsystem CFD modeling: taking a transonic compressor single-stage Rotor NASA Rotor67 as a diagnostic sample, performing static/dynamic aeroelastic solution on aeroelastic response by a flow-solid coupling numerical simulation method, and firstly completing a subsystem: solving a model of a flow field and a structure;
S2, modeling by a subsystem CSD: determining a structural dynamic response by discretizing a Finite Element Analysis (FEA) on the spatial variable and implementing a time solution of the time variable of the global aeroelastic motion equation;
S3, static/dynamic gas-elastic integrated solving model: carrying out integrated solution on the flow-solid coupling algorithm and the correction blade profile transmission to realize accurate solution of the thermal state shape;
S4, establishing a damaged blade parameterized model: based on the thermal profile, introducing damage parameters to simulate potential cracks or damage within the blade structure;
s5, vibration signal analysis of the aerodynamic elastic response: analyzing damage parameters of damaged blades through three-dimensional damage positions h d, damage ranges l d and damage grades Sr to obtain detection vibration forms, vibration orders and attenuation coefficients, and using the data to diagnose damage conditions of the blades;
s6, diagnosing the damage position, the damage range and the damage degree: through the combination and comparison of various characteristics, the vibration form, the frequency spectrum information and the pneumatic elasticity stability are comprehensively considered, and the coupling analysis is carried out so as to comprehensively diagnose the position, the range and the degree of the damage.
Preferably, subsystem CFD modeling in S1 uses a solution to the three-dimensional reynolds average Navier-Stokes (RANS) equation under a rotating coordinate system in a steady flow field simulation (formula (1)). The rotational angular velocity vector is ω= [ ω,0] T. The control equation is discretized by adopting a finite volume method:
Wherein Ω represents a fixed control volume having a boundary S, and the solution vector W, the convection flux F c, the viscous flux F v, and the source term Q can be written as component forms shown in formula (2), respectively, σ ij is given by formula (3):
where ρ represents density, p represents pressure, E represents the relative total energy per unit mass, K represents thermal conductivity, and T represents temperature.
Preferably, the time solution of the time variable of the Finite Element Analysis (FEA) discrete on the space variable and realizing the global aeroelastic motion equation in the S2 is specifically that
Where M represents the system mass, C represents the damping, K represents the stiffness matrix, F is the external load vector of the system node obtained by integrating the respective identity matrix and vector,Representing acceleration of the structure,/>Representing velocity, u represents displacement.
Preferably, the numerical integration method for solving the global aeroelastic motion equation relies on a basic assumption that the displacement, velocity and acceleration maintain a simplified linear relationship during each time interval Δt, and the dynamic response analysis proceeds using the Newmark Beta method by monitoring the time domain displacement response of the structure after a small disturbance is applied, the response being calculated at these discrete time points.
[A1]{un+1}＝[A2]+[A3]{un}+[A4]{un-1} (5)
Wherein, the algorithm of each parameter is as follows:
[A1]＝[M/Δt2+K/3]
[A3]＝[2M/Δt2-K/3]
[A4]＝[-M/Δt2-K/3] (6)
the convergence of the displacement increment determines whether the iteration is continued or not, and when the calculated displacement increment is sufficiently small, no significant influence is generated on the calculated result, so that the method is considered to be close enough to an accurate solution, namely, the real solution is approximated by continuous iteration.
Preferably, the step S3 is to perform unsteady flow field solution through CFD/CSD coupling time domain propulsion of a dynamic aeroelastic solution module so as to obtain a static deformation result strongly coupled with the flow field, thereby realizing accurate solution of the thermal state shape.
Preferably, the static/dynamic pneumatic elastic integrated solving flow is as follows
① Under a given calculation working condition, inputting the cold state (processing) appearance and the initial grid of the blade into a CFD/CSD coupling module;
② Performing time domain unsteady propulsion on the coupling module until the solution converges, wherein the output thermal state (working) appearance is the solution result of the static pneumatic elastic module;
③ The thermal state appearance is used as a correction blade type to be input into the dynamic aeroelastic module for solving;
④ Performing CFD/CSD time domain propulsion after applying small disturbance deformation;
⑤ And analyzing the pneumatic elastic response behavior of the blade according to the time domain response curve of the blade displacement.
Preferably, the position of the damage in the S4 has equivalent rigidity degradation effect on the surrounding positions, and a rigidity loss coefficient is definedThus quantifying the extent of the lesions, two other dimensionless lesion parameters of l d = H/H and H d＝h0/H are defined, characterizing the extent and location of the lesions, respectively, with the possible lesions being located at H d of length l d.
Preferably, in the step S5, the vibration form is that the displacement response of the damaged blade time domain is obtained through high-fidelity CFD/CSD coupling solution; selecting characteristic points and recording vibration signals; analyzing the vibration data to identify the vibration form of the blade, and comprehensively investigating the linear and nonlinear effects; in addition, the amplitude of the vibration directly affects the fatigue life of the structure.
Preferably, the vibration order in S5 is to convert the vibration signal into frequency domain data by using a Fast Fourier Transform (FFT) tool; determining main vibration frequency components of the blade according to the frequency spectrum analysis result; the frequency components are monitored for changes with critical parameters of the lesion to identify the location, extent, or extent of the lesion.
Preferably, the damping coefficient in S5 is an exponential function used to evaluate the change of vibration amplitude with time by envelope analysis of the vibration signalFitting characteristic point displacement response curve peak values,/>The aeroelastic stability of the structure is characterized for the attenuation coefficient.
The invention provides a method for diagnosing the damage position of a rotor of a transonic compressor by aeroelastic analysis, which has the following beneficial effects compared with the prior art:
1. the parameter changes of the damage degree, the damage range and the damage position can be effectively and sensitively reflected through vibration information analysis;
2. in the aspect of aeroelastic stability, the frequency spectrum information and attenuation coefficient of vibration are proved to be sensitive to the change of the damage position, the damage range and the damage degree, and can be used as an effective tool for detecting the structural damage;
3. compared with general damage monitoring, the method can judge the positive/negative influence of different damage parameters on the stability of the gas bomb, and further predict potential dangerous damage.
Drawings
FIG. 1 is a flow chart of the diagnostic method of the present invention.
Fig. 2 is a physical diagram and numerical values of a transonic single stage compressor Rotor NASA Rotor 67.
Fig. 3 is a numerical simulation of a transonic single stage compressor Rotor NASA Rotor 67.
Fig. 4 is a static/dynamic pneumatic elastic integrated solution flowchart in the invention S3.
Fig. 5 is a parametric modeling diagram of damaged blades in the invention S4.
FIG. 6 is a graph of the vibration damping coefficient definition and the working line damping coefficient of a healthy blade near stall point in example 1 of the present invention.
FIG. 7 is a plot of spanwise and chordwise stress distribution at the near stall point vibration extremum for a healthy blade according to example 1 of the present invention.
FIG. 8 is a graph showing the time domain response and attenuation coefficient of the damage range 0.3 in example 1 of the present invention.
FIG. 9 is a graph showing the time domain response and attenuation coefficient of the damage range 0.5 in example 1 of the present invention.
FIG. 10 is a graph showing the time domain response and attenuation coefficient of the damage range 0.7 in example 1 of the present invention.
FIG. 11 is a graph of FFT versus attenuation coefficient for the range of 0.2 for example 2 of the present invention.
FIG. 12 is a graph of FFT versus attenuation coefficient for the range of 0.4 for example 2 according to the present invention.
FIG. 13 is a graph of FFT versus attenuation coefficient for the range of 0.2 for example 3 according to the present invention.
FIG. 14 is a graph showing the limit cycle oscillation of the damage level 0.1 and the damage range 0.2 in example 3 of the present invention.
Detailed Description
The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention, and it is apparent that the described embodiments are only some embodiments of the present invention, but not all embodiments, and all other embodiments obtained by those skilled in the art without making creative efforts based on the embodiments of the present invention are included in the protection scope of the present invention.
Referring to fig. 1-14, the embodiment of the present invention provides a technical solution: the method for diagnosing the damage position of the rotor of the transonic compressor by aeroelastic analysis specifically comprises the following steps:
The subject of the present invention (as shown in fig. 2) is a transonic compressor single stage Rotor NASA Rotor67. Wherein the damaged blade is modeled by three dimensional parameterization: lesion location h d, lesion extent l d, and lesion level Sr. The aeroelastic response was solved for static/aerodynamic elasticity by a flow-solid coupled numerical simulation method (as shown in fig. 3).
S1, subsystem CFD modeling:
For the illustrated transonic single stage compressor Rotor NASA Rotor67, the blade will exhibit a aeroelastic response under the coupling of aerodynamic forces, elastic forces and inertial forces, which is caused by the interaction of the structural and aerodynamic modules, so the primary step in building the aeroelastic model is to complete the subsystem: and solving a model of a flow field and a structure.
In the unsteady flow field simulation adopted by the invention, a three-dimensional Reynolds average Navier-Stokes (RANS) equation under a rotating coordinate system is solved (formula (1)). The rotational angular velocity vector is ω= [ ω,0] T. The control equation described above is discretized using a finite volume method.
Where Ω denotes a fixed control volume having a boundary S, the solution vector W, the convection flux F c, the viscous flux F v, and the source term Q can be written as component forms shown in formula (2), respectively, and σ ij is given by formula (3).
Wherein ρ, p, E, K, T represent density, pressure, relative total energy per unit mass, thermal conductivity, and temperature, respectively. Further, r i、ui and ω i refer to displacement, velocity, and speed in three directions relative to a rotating reference frame. σ ij is the viscous stress tensor with dynamic viscosity coefficient μ, which can be approximated as an ideal gas by the Sutherland law. The discrete equation adopts a second-order backward Euler technique to carry out time integration, and the space discrete adopts a second-order upwind method. All simulations were solved using double precision, with the turbulence model employed being the SST k- ω model. The stagnation parameters and the angle of inflow are given at the inlet and the radial pressure balance conditions are specified at the outlet. In reaching the stall point, a reliable boundary is obtained by setting a mass flow boundary at the outlet.
S2, modeling by a subsystem CSD:
the structural dynamic response may be determined by discretizing a Finite Element Analysis (FEA) on the spatial variables and implementing a time solution to the time variable of the global aeroelastic equation of motion.
Wherein M, C, K represent the system mass, damping and stiffness matrices, respectively, and F is the external load vector of the system node, obtained by integrating the respective identity matrices and vectors.And u is the acceleration, velocity and displacement of the structure. C includes inherent structural damping and rotation induced coriolis force effects with limited impact, and therefore this value is typically ignored for greater safety design margin. The numerical integration method used to solve the equation relies on a basic assumption that the displacement, velocity and acceleration remain in a simplified linear relationship for each time interval Δt. The response is calculated at these discrete time points, advanced using the Newmark-Beta method.
[A1]{un+1}＝[A2]+[A3]{un}+[A4]{un-1} (5)
Wherein,
[A1]＝[M/Δt2+K/3]
[A3]＝[2M/Δt2-K/3]
[A4]＝[-M/Δt2-K/3] (6)
The convergence of the displacement increment determines whether iteration is continued or not, when the calculated displacement increment is small enough, no obvious influence is generated on the calculated result, so that the method is considered to be close to an accurate solution, namely, the true solution is approximated through continuous iteration, and the reliability of the calculated result is improved. The loads on the blade include inertial and aerodynamic loads. Dynamic response analysis is then performed by monitoring the time domain displacement response of the structure after a small disturbance has been applied.
S3, static/dynamic gas-elastic integrated solving model:
blade aerostatic elasticity involves a solution of static deformation under given flow conditions, typically implemented using Finite Element Methods (FEM) or Boundary Element Methods (BEM), in which the interplay between the blade and its detour is ignored. In order to ensure that the blade can reach the designed appearance and performance index under the working state, in the aeroelastic numerical simulation to be developed, static aeroelastic is not solved based on the traditional steady/steady aerodynamic force, but rather the unsteady flow field is solved through CFD/CSD coupling time domain propulsion of a dynamic aeroelastic solving module, so as to obtain the static deformation result strongly coupled with the flow field, and the accurate solving of the 'thermal state' appearance (appearance under the working rotating speed) is realized, so that the geometrical nonlinear effect caused by the rotation of the blade is comprehensively considered, and the method comprises the following steps: ① The rotating blades are subjected to an outward centrifugal force (tensile force) resulting in a stress stiffening effect that significantly increases their own lateral stiffness; ② The rotation of the blades results in a rotational softening effect which reduces the stiffness of the blade itself in a direction other than perpendicular to the plane of rotation; ③ The effect of the following forces present between centrifugal force and blade displacement.
On the other hand, the thermal state appearance obtained by static aeroelasticity is used as an initial structure to be input into a dynamic aeroelastic solving process for carrying out leaf type correction, and because the unsteady solving process generally needs to carry out longer time domain pushing to achieve a balanced state between load and deformation, the input of the thermal state appearance avoids repeated calculation of deformation, saves calculation resources and participates in the solving of dynamic response of the structure in an appearance which is closer to an actual working state.
The specific calculation flow is shown in fig. 4, and the solving process is as follows:
① Under a given calculation working condition, inputting the cold state (processing) appearance and the initial grid of the blade into a CFD/CSD coupling module;
② Performing time domain unsteady propulsion on the coupling module until the solution converges, wherein the output thermal state (working) appearance is the solution result of the static pneumatic elastic module;
③ The thermal state appearance is used as a correction blade type to be input into the dynamic aeroelastic module for solving; ④ Performing CFD/CSD time domain propulsion after applying small disturbance deformation;
⑤ And analyzing the pneumatic elastic response behavior of the blade according to the time domain response curve of the blade displacement.
In summary, the static-dynamic pneumatic elasticity is integrated and solved with the correction blade profile transmission through a flow-solid coupling algorithm, so that the solving precision, efficiency and reliability are improved.
S4, establishing a damaged blade parameterized model:
The location, extent and extent of material degradation of the damage may negatively impact the aeroelastic response of the blade, and therefore the present invention introduces damage parameters to simulate potential cracks or damage in the blade structure, and high speed transonic compressor rotor blades tend to reach equilibrium conditions of thermal profiles in a very short time, and therefore the parametric basis model of damaged blades of the present invention is thermal profiles, as shown in FIG. 5, the location of the damage is considered to produce an equivalent stiffness degradation effect on its surrounding locations, and thus the stiffness loss coefficient is defined The extent of damage was quantified. It should be noted that the smaller the Sr, the greater the extent of damage. Furthermore, the possible lesions are located at H d of length l d, and the extent and location of the lesions are characterized by defining two further dimensionless lesion parameters of l d =h/H and H d＝h0/H, respectively.
It should be noted that the damage model is constructed by accounting for degradation of local material properties and does not change the existing geometry of the structure. In addition, the main object of the present invention is to study the effect of damage on structural vibration response and not to the process of damage occurrence.
S5, vibration signal analysis of the aerodynamic elastic response:
Vibration data of damaged blades can be collected through pneumatic and pneumatic elastic solving, damage parameters of the damaged blades are analyzed through three dimensions, and the damage parameters are obtained:
① Vibration form: obtaining displacement response of a damaged blade time domain through high-fidelity CFD/CSD coupling solution; selecting characteristic points and recording vibration signals; analyzing the vibration data to identify the vibration form of the blade, and comprehensively investigating the linear and nonlinear effects; in addition, the amplitude of the vibration directly affects the fatigue life of the structure.
② Vibration order: converting the vibration signal into frequency domain data using a Fast Fourier Transform (FFT) tool; determining main vibration frequency components of the blade according to the frequency spectrum analysis result; the frequency components are monitored for changes with critical parameters of the lesion to identify the location, extent, or extent of the lesion.
③ Attenuation coefficient: using envelope analysis of vibration signals to evaluate vibration amplitude over time, using exponential functionsFitting characteristic point displacement response curve peak values,/>The aeroelastic stability of the structure is characterized for the attenuation coefficient.
By these means, the data can be used to diagnose damage to the blade by monitoring vibration, frequency and attenuation characteristics of the blade in real time. The monitoring method adopted by the invention can help predict the potential blade damage problem so as to take timely maintenance and repair measures and ensure the performance and safety of the equipment.
S6, diagnosing the damage position, the damage range and the damage degree:
The vibration form, the frequency spectrum information and the pneumatic elasticity stability are comprehensively considered, and the coupling analysis is carried out to comprehensively diagnose the position, the range and the degree of damage, and the combination and the comparison of various characteristics are involved so as to provide the most accurate diagnosis result.
① And (3) data acquisition: vibration data and spectrum information are collected in the dynamic response process of the damaged blade, and information such as the amplitude, frequency, vibration mode, spectrum distribution and the like of the vibration of the blade is captured.
② Spectral feature extraction: for vibration data and spectrum information, characteristics related to damage are extracted, including changes in vibration modes, movements and changes in spectrum peaks, anomalies in spectrum distribution, and changes in aeroelastic stability.
③ Determination of vibration form and damping coefficient: and obtaining an attenuation coefficient and a variation trend thereof through dynamic response data, and determining attenuation and divergence of vibration or limit cycle response with constant amplitude.
④ Evaluation of the injury position: the damage position is sensitive to the vibration form and the amplitude range of the frequency spectrum, and the larger frequency spectrum amplitude corresponds to the damage starting position which is closer to the root of the blade; the damage position is sensitive to the change trend of the attenuation coefficient, and the damage position close to the root causes the change trend of the attenuation coefficient to be more complex;
⑤ Evaluation of injury scope: the range of damage is sensitive to the shift of the spectral peaks, which will tend to shift to lower frequencies as the range of damage increases. A larger lesion field may result in a significant shift in spectral peaks, which is directly related to the location and extent of the lesion.
⑥ Evaluation of injury level: the impairment level is more sensitive to the frequency information of the spectrum, and higher impairment levels will lead to a significant increase in the high frequency content of the spectrum. In addition, the damage level directly causes a change in the attenuation coefficient of the structure.
And all that is not described in detail in this specification is well known to those skilled in the art.
For a transonic compressor single-stage Rotor blade NASA Rotor67, simulation parameters are specifically shown in table 1, the materials are selected to be titanium alloys, the density is 4440kg/m < 3 >, the Young's modulus is 142GPa, and the Poisson's ratio is 0.3. And selecting the front edge point of the blade tip as a characteristic point to acquire vibration information, wherein the plotted graph is the response of the characteristic point if no special description exists.
According to the flow-solid coupled dynamic response solving result of the healthy blade, as shown in fig. 6, the attenuation coefficient of the healthy blade is given by vibration information, and the NASA Rotor67 has no risk of chatter in the working line range. The characteristics of displacement distribution, equivalent stress, strain distribution and the like at the vibration extreme points are examined, and the phenomenon of stress concentration is found in the part of the structure from the leaf height of 30% of the spanwise distance casing to the root and from the chordwise distance front edge to 1/3 to 2/3, as shown in fig. 7. This will result in a significant increase in stress, possibly exceeding the fatigue limit of the material, thereby rapidly causing fatigue damage to the material, while these stress concentration areas often become the initiation sites for cracks, which after multiple loads may propagate in these areas, eventually leading to fatigue fracture of the material. Thus, the lesion location h d = 0.3,0.6,0.8 was selected and applied and analyzed as an example for the extent and extent of lesions that are not feasible.
Table 1 pneumatic design parameters
| Rotational speed | 16043rpm |
| Tip speed | 429m/s |
| Blade tip relative Mach number | 1.38 |
| Aspect ratio | 1.56 |
| Number of blades | 22 |
| Blade tip clearance | 1.016mm |
Example 1: the lesion position h d =0.3, the lesion degree sr= 0.2,0.4,0.6,0.8,1, the lesion range l d = 0.3,0.5,0.7
As shown in fig. 8 to 10, when the damage start position appears at a position 30% away from the blade tip, the damage range is selected from the middle part of the blade gradually to the root part of the blade, and the damage degree is from that of a healthy blade (sr=1) to that of a damaged blade (sr=0.2) with 80%, and the vibration signals under different working conditions show obvious differences.
The influence of Sr is transversely compared, the increase of the damage degree directly causes the increase of the vibration amplitude and the backward shift of the vibration phase, the amplitude is sensitive to Sr in the initial stage of vibration, but the positive attenuation coefficient enables the amplitude under each parameter to be attenuated to the same level quickly, and the sensitivity is reduced.
Attenuation coefficientCharacterization of the aeroelastic stability of the injury system shows that the change in the extent of injury and the extent of injury have a significant impact on their stability. Unlike expectations, the increase in the degree of damage does not merely result in a decrease in the structural aeroelastic stability, but rather has a positive effect in a certain range, where this boundary is referred to as "critical" Sr, it being noted that "critical" refers not to the critical of stability and destabilization, but to the critical of increased stability and reduced stability. Obviously, at this lesion position h d =0.3, the lesion parameter is l d =0.7, sr=0.2 being the least stable.
The results show that when the damage position is fixed, the aeroelastic stability is very sensitive to the damage degree Sr, and the change of the damage parameters in the aspect can directly cause the change of the attenuation coefficient and the frequency domain information.
Example 2: the injury position h d =0.6, the injury degree sr= 0.1,0.3,0.5,0.7,0.9,1, the injury range l d =0.2, 0.4
To examine the different damage patterns, the damage start position was shifted down to 60% from the tip, and the damage level was selected from the characteristic point response of healthy blade (sr=1) until 90% of damaged blade (sr=0.1) occurred.
Fig. 11 to 12 respectively show the damage range of h d =0.6, the damage range of l d =0.2, and the spectrum information and attenuation coefficient information of different damage degrees under 0.4, and the result shows that the spectrum signal is very sensitive to the change of the damage degree, especially in a high frequency region, through Fast Fourier Transform (FFT) and transverse comparison and fixation of different Sr working conditions of l d. Longitudinally comparing different l d working conditions, it can be found that a larger l d will result in an FFT graph with more amplitude separation. In particular for a blade damaged by h d＝0.6,ld =0.4, sr=0.1, where the damping coefficient is negative, i.e. the vibrations are no longer damped, the amplitude gradually increases, the spectral signal contains more complex vibration information.
From the angle of the attenuation coefficient, it can be found that different damage positions cause obvious difference of each working condition on critical Sr by longitudinally comparing with example 1, the damage starting position of example 1 is positioned at 30% of the blade, the attenuation coefficient of the example 1 shows a rule of increasing firstly and then decreasing with the decrease of Sr (the increase of the damage degree), and the damage position of example 2 is positioned at 60% of the blade, and decreases firstly and then increases and then decreases with the increase of the damage degree.
In summary, the damage range and the damage degree are sensitive to the spectrum signal, and the damage position is sensitive to the change rule of the attenuation coefficient.
Example 3: the lesion position h d =0.8, the lesion degree sr= 0.1,0.3,0.5,0.7,0.9,1, and the lesion range l d =0.2
The damage start position continues to move down to near the root, i.e. 80% from the tip, the extent of damage is selected from the characteristic point response of a healthy blade (sr=1) until 90% of the damaged blade (sr=0.1) has occurred.
Fig. 13 to 14 show the FFT, attenuation coefficient, and nonlinear response results of h d＝0.8,ld =0.2, sr= 0.1,0.3,0.5,0.7,0.9,1, respectively. The results indicate that the damage occurring at the root is more complex, but if the propagation in the spanwise direction is not continued, there is some nonlinear effect on the structure itself so that the damaged blade amplitude does not diverge, i.e. exhibits nonlinear limit cycle oscillations (fig. 14). In addition, the degree of damage is sensitive to both the amplitude and frequency of the vibration spectrum information, and the attenuation coefficient also shows a change law similar to that of example 2.
In summary, the vibration form, the spectrum information of the vibration signal, and the aeroelastic stability (i.e., attenuation coefficient) are sensitive to the damage position, the damage range, and the damage degree, and can be used as a tool for diagnosing the damage position and the damage early warning, so as to predict the fatigue life through the amplitude and the frequency.
The invention starts from the real response process of the compressor rotor, integrates static/dynamic pneumatic elastic solving, replaces static solving with unsteady pushing to improve the solving precision of the thermal state (under the working rotating speed) appearance, uses the thermal state appearance as a reference model of dynamic pneumatic elastic solving to further improve the solving precision of the dynamic pneumatic elasticity, carries out parameterization modeling on damaged blades on the basis, and carries out diagnosis on the damage position, the damage range and even the damage degree through analysis of the pneumatic elastic response characteristic of the damaged blades.
The results show that different damage parameters will cause the vibration to exhibit a significant frequency shift and typical nonlinear response form-limit cycle. Compared with the classical vibration-based method, the method has higher sensitivity, can simultaneously consider the influence of the damage degree, the damage range and the damage position, and can predict the damage parameters of potential aeroelastic instability on the basis so as to early warn in advance.
It is noted that relational terms such as first and second, and the like, are used solely to distinguish one entity or action from another entity or action without necessarily requiring or implying any actual such relationship or order between such entities or actions, and further, that the terms "comprise," "include," or any other variation thereof, are intended to cover a non-exclusive inclusion such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements, but may include other elements not expressly listed or inherent to such process, method, article, or apparatus.
Although embodiments of the present invention have been shown and described, it will be understood by those skilled in the art that various changes, modifications, substitutions and alterations can be made therein without departing from the principles and spirit of the invention, the scope of which is defined in the appended claims and their equivalents.
Claims (10)
1. The method for diagnosing the damage position of the rotor of the transonic compressor by aeroelastic analysis is characterized by comprising the following steps of:
S1, subsystem CFD modeling: taking a transonic compressor single-stage Rotor NASA Rotor67 as a diagnostic sample, performing static/dynamic aeroelastic solution on aeroelastic response by a flow-solid coupling numerical simulation method, and firstly completing a subsystem: solving a model of a flow field and a structure;
S2, modeling by a subsystem CSD: determining a structural dynamic response by discretizing a Finite Element Analysis (FEA) on the spatial variable and implementing a time solution of the time variable of the global aeroelastic motion equation;
S3, static/dynamic gas-elastic integrated solving model: carrying out integrated solution on the flow-solid coupling algorithm and the correction blade profile transmission to realize accurate solution of the thermal state shape;
S4, establishing a damaged blade parameterized model: based on the thermal profile, introducing damage parameters to simulate potential cracks or damage within the blade structure;
s5, vibration signal analysis of the aerodynamic elastic response: analyzing damage parameters of damaged blades through three-dimensional damage positions h d, damage ranges l d and damage grades Sr to obtain detection vibration forms, vibration orders and attenuation coefficients, and using the data to diagnose damage conditions of the blades;
s6, diagnosing the damage position, the damage range and the damage degree: through the combination and comparison of various characteristics, the vibration form, the frequency spectrum information and the pneumatic elasticity stability are comprehensively considered, and the coupling analysis is carried out so as to comprehensively diagnose the position, the range and the degree of the damage.
2. The method for diagnosing a damaged position of a rotor of a transonic compressor according to claim 1, wherein the subsystem CFD modeling in S1 uses a non-stationary flow field simulation by solving a three-dimensional reynolds average Nav er-Stokes (RANS) equation in a rotation coordinate system (formula (1)), the rotation angular velocity vector is ω= [ ω,0] T, and the control equation is discretized by a finite volume method:
Wherein Ω represents a fixed control volume having a boundary S, and the solution vector W, the convection flux F c, the viscous flux F v, and the source term Q can be written as component forms shown in formula (2), respectively, σ ij is given by formula (3):
where ρ represents density, p represents pressure, E represents the relative total energy per unit mass, K represents thermal conductivity, and T represents temperature.
3. The method for diagnosing a damaged position of a rotor of a transonic compressor according to claim 1, wherein the time solution of the time variable of the Finite Element Analysis (FEA) discrete on the space variable and implementing the global aeroelastic motion equation in S2 is specifically:
where M represents the system mass, C represents the damping, K represents the stiffness matrix, F is the external load vector of the system node obtained by integrating the respective identity matrix and vector, Representing acceleration of the structure,/>Representing velocity, u represents displacement.
4. A method for diagnosing a damaged position of a rotor of a transonic compressor by aeroelastic analysis according to claim 3, characterized in that the numerical integration method for solving the global aeroelastic equation of motion relies on a basic assumption that the displacement, velocity and acceleration remain in a simplified linear relationship during each time interval Δt, and the dynamic response analysis is performed by monitoring the time domain displacement response of the structure after applying small perturbations, the response being calculated at these discrete time points, using the Newmark Beta method for the hypothetical propulsion:
[A1]{un+1}＝[A2]+[A3]{un}+[A4]{un-1} (5)
wherein, the algorithm of each parameter is as follows:
the convergence of the displacement increment determines whether the iteration is continued or not, and when the calculated displacement increment is sufficiently small, no significant influence is generated on the calculated result, so that the method is considered to be close enough to an accurate solution, namely, the real solution is approximated by continuous iteration.
5. The method for diagnosing a damaged position of a rotor of a transonic compressor according to claim 1, wherein the step S3 is to solve an unsteady flow field by CFD/CSD coupled time domain propulsion of a dynamic gas-elastic solving module so as to obtain a static deformation result strongly coupled with the flow field, thereby realizing accurate solving of a thermal state shape.
6. The method for diagnosing a damaged position of a rotor of a transonic compressor according to claim 5, wherein the static/dynamic pneumatic elastic integrated solution flow is as follows:
① Under a given calculation working condition, inputting the cold state (processing) appearance and the initial grid of the blade into a CFD/CSD coupling module;
② Performing time domain unsteady propulsion on the coupling module until the solution converges, wherein the output thermal state (working) appearance is the solution result of the static pneumatic elastic module;
③ The thermal state appearance is used as a correction blade type to be input into the dynamic aeroelastic module for solving;
④ Performing CFD/CSD time domain propulsion after applying small disturbance deformation;
⑤ And analyzing the pneumatic elastic response behavior of the blade according to the time domain response curve of the blade displacement.
7. The method for diagnosing a damaged position of a rotor of a transonic compressor according to claim 1, wherein the damaged position in S4 has an equivalent stiffness degradation effect on surrounding positions, and a stiffness loss coefficient is definedThus quantifying the extent of damage, the smaller Sr, the greater the extent of damage, and furthermore, the damage is located at H d of length L d = H/H, defining dimensionless damage parameters of L d＝ld/H and H d＝h0/H, where L d and H d characterize the extent and location of the damage, respectively.
8. The method for diagnosing a damaged position of a rotor of a transonic compressor according to claim 1, wherein the vibration form in S5 is a displacement response of a damaged blade in a time domain obtained by high-fidelity CFD/CSD coupling solution; selecting characteristic points and recording vibration signals; analyzing the vibration data to identify the vibration form of the blade, and comprehensively investigating the linear and nonlinear effects; in addition, the amplitude of the vibration directly affects the fatigue life of the structure.
9. The method for diagnosing a damaged position of a rotor of a transonic compressor according to claim 1, wherein the vibration order in S5 is obtained by converting a vibration signal into frequency domain data using a Fast Fourier Transform (FFT) tool; determining main vibration frequency components of the blade according to the frequency spectrum analysis result; the frequency components are monitored for changes with critical parameters of the lesion to identify the location, extent, or extent of the lesion.
10. The method for diagnosing a damaged position of a rotor of a transonic compressor according to claim 1, wherein the damping coefficient in S5 is an exponential function using envelope analysis of vibration signals to evaluate the change of vibration amplitude with timeFitting characteristic point displacement response curve peak values,/>The aeroelastic stability of the structure is characterized for the attenuation coefficient.
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