A Numerical Analysis Method for Aerodynamic Flutter of Turbine Blade Thermal Barrier Coatings at High Temperatures

Abstract

The present invention discloses a numerical analysis method for gas flow excitation of a thermal barrier coating on a turbine blade at high temperatures. An external flow field geometric model and a geometric model of a turbine blade with a thermal barrier coating are established, and computational grids are respectively set; the finite element software is used to perform modal analysis on the geometric model of the turbine blade to obtain its resonance characteristics and critical speed; the fluid analysis software is used to perform steady-state aerodynamic analysis on the external flow field geometric model to obtain its steady-state flow field and steady-state temperature field, and the steady-state flow field and steady-state temperature field are used as the initial conditions of the unsteady flow field for unsteady aerodynamic analysis to obtain the pressure field on the surface of the turbine blade that changes periodically with time; the obtained pressure field is interpolated onto the surface of the thermal barrier coating, the steady-state temperature field is imported into the thermal barrier coating and the turbine blade, and transient response analysis is performed to obtain the stress field and displacement field at different times; the fast Fourier transform is used to perform time-frequency spectrum transformation on the pressure field, stress field, and displacement field data at the dangerous position points to obtain the dynamic response amplitude and frequency of the thermal barrier coating under gas flow excitation.

Description

Technical Field

[0001] The invention belongs to the technical field of performance analysis of thermal insulation protective coatings of aircraft engines, and in particular relates to a numerical analysis method for airflow excitation of thermal barrier coatings of turbine blades under high temperature. Background Art

[0002] Thermal barrier coatings (TBCs) are a layer of ceramic coatings deposited on the surface of high-temperature resistant metals or superalloys. Thermal barrier coatings act as a heat insulator for the substrate material, which can reduce the substrate temperature so that the devices made of them (such as engine turbine blades) can operate at high temperatures. They have the characteristics of high melting point, low thermal conductivity, corrosion resistance, and thermal shock resistance. During high-temperature service, thermal barrier coatings can protect high-temperature substrates, increase the temperature and thermal efficiency of thermal engines, and are therefore widely used in the fields of aviation, chemical industry, metallurgy, and energy.

[0003] However, in the actual application process, due to the mismatch of material parameters and thermal residual stress, high-temperature sintering effect of ceramic materials, high-temperature interface oxidation, etc., cracks are prone to appear inside the coating, and the thermal barrier coating of the turbine blade vibrates at high frequency under the wake of the upstream stator, and the cracks inside the coating will quickly expand and peel off under high-frequency vibration. Once the coating peels off, the base metal parts will be exposed to high temperature environment, and the consequences are very serious.

[0004] Therefore, it is very important to study the vibration stress of thermal barrier coatings under airflow excitation, which can not only be used to analyze the failure process and causes of thermal barrier coating spalling, but also be used for coating design to extend the service life of the coating after microcracks occur. However, there is no relatively mature numerical simulation method for analyzing the airflow excitation force of thermal barrier coatings in the prior art. Summary of the invention

[0005] In order to overcome the shortcomings of the above-mentioned prior art, the purpose of the present invention is to provide a numerical analysis method for airflow excitation of turbine blade thermal barrier coating at high temperature, so as to numerically simulate the airflow excitation force of turbine blade thermal barrier coating at high temperature, and provide a reference for high cycle fatigue failure of thermal barrier coating and coating design.

[0006] In order to achieve the above object, the technical solution adopted by the present invention is:

[0007] A numerical analysis method for airflow excitation of a turbine blade thermal barrier coating under high temperature comprises the following steps:

[0008] Step 1, establishing an external flow field geometry model and a turbine blade geometry model with a thermal barrier coating, and setting calculation grids respectively;

[0009] Step 2, using finite element software to perform modal analysis on the geometric model of the turbine blade containing the thermal barrier coating to obtain its resonance characteristics and critical speed;

[0010] Step 3, using fluid analysis software to perform steady-state aerodynamic analysis on the external flow field geometric model to obtain its steady-state flow field and steady-state temperature field; using the obtained steady-state flow field as the initial condition of the unsteady-state flow field, performing unsteady-state aerodynamic analysis on the external flow field geometric model to obtain the pressure field on the turbine blade surface that changes with time period;

[0011] Step 4, interpolating the pressure field that changes with time periodically to the surface of the thermal barrier coating in the geometric model of the turbine blade containing the thermal barrier coating, importing the steady-state temperature field into the thermal barrier coating and the turbine blade, performing transient response analysis, and obtaining stress fields and displacement fields at different times;

[0012] Step 5, using fast Fourier transform, the pressure field, stress field and displacement field data of the dangerous position point are transformed into time spectrum to obtain the dynamic response amplitude and frequency of the thermal barrier coating under airflow excitation. The dangerous position point refers to the position where the stress is greater than the set value.

[0013] In one embodiment, the step 1 is to establish an external flow field geometry model and a turbine blade geometry model containing a thermal barrier coating in a finite element modeling software, wherein the external flow field geometry model is a periodic single-channel fluid geometry model; and to establish a solid domain computational grid and a periodic fluid domain computational grid of the thermal barrier coating of the turbine blade in a meshing software.

[0014] In one embodiment, the external flow field geometric model and the turbine blade geometric model containing the thermal barrier coating are established, and the implementation method is as follows:

[0015] Step 101, respectively establishing a turbine stationary blade geometric model and a turbine moving blade geometric model;

[0016] Step 102, establishing a thermal barrier coating on the turbine blade substrate of the turbine blade geometric model to obtain a turbine blade geometric model containing the thermal barrier coating; the turbine stator blade geometric model and the turbine blade geometric model containing the thermal barrier coating constitute the turbine blade geometric model containing the thermal barrier coating;

[0017] Step 103, respectively establish a turbine stator flow field geometry model and a turbine rotor flow field geometry model, and connect the geometric positions of the two flow field geometry models to form the external flow field geometry model, and the external flow field is a unit channel of a periodic turbine blade.

[0018] In one embodiment, the step 1 is to establish a solid domain computational grid for the thermal barrier coating of the turbine blade, and the implementation method is as follows:

[0019] Dividing the geometric model of the turbine stator blade and the geometric model of the turbine moving blade with the thermal barrier coating into meshes, naming the mesh parts of the geometric model of the turbine moving blade with the thermal barrier coating as the blade root part, the blade tip part and the coating part, and exporting the blade root part, the blade tip part and the coating part as well as the interface between the blade base and the thermal barrier coating into the mesh format of the finite element software used;

[0020] The periodic fluid domain computational grid is established as follows:

[0021] The turbine stator flow field geometry model and the turbine moving blade flow field geometry model are meshed, and the boundary layer mesh is divided on the blade surface. The mesh inlet, outlet, blade surface, and periodic interface are named respectively and exported to the mesh format of the fluid calculation software.

[0022] In one embodiment, the step 2 is implemented as follows:

[0023] Step 201, importing the solid domain calculation grid of the turbine blade thermal barrier coating into finite element software, and defining the material parameters of the blade substrate and the thermal barrier coating respectively, including density, elastic modulus, Poisson's ratio, thermal conductivity and specific heat capacity;

[0024] Step 202, defining boundary conditions of a geometric model of a turbine blade with a thermal barrier coating, defining fixed boundary conditions, elastic support boundary conditions and contact boundary constraint conditions for the root portion of the turbine blade, and defining free surfaces for other portions of the turbine blade;

[0025] Step 203, after setting the number of modes to be solved and the rotation speed, the natural frequency of the thermal barrier coating of the turbine blade at different rotation speeds, i.e., the Campbell diagram, is calculated, the critical rotation speed ω at which resonance occurs is determined, and the coordinate value of the point to be analyzed is found from the relative stress and displacement of the thermal barrier coating under resonance.

[0026] In one embodiment, the fixed boundary condition is that at the fixed point, the displacement in all directions is zero; the elastic support boundary condition is that a part of the structure is constrained by elastic support, and the displacement of the structure at the support point is proportional to the reaction force received; the contact boundary constraint condition is the interaction between two or more parts of the structure. In the contact area, the structural parts can separate or slide but cannot penetrate each other.

[0027] In one embodiment, the step 3 is implemented as follows:

[0028] Step 301, import the turbine stator flow field geometry model and turbine moving blade flow field geometry model of the divided calculation grid into the fluid analysis software, define the material of the fluid calculation domain as ideal air, adopt the turbulence model and the non-equilibrium near-wall model, set the inlet static pressure, temperature, turbulence and outlet static pressure, adopt the frozen rotor method for calculation, set the iterative step solution, and obtain the steady-state flow field and steady-state temperature field after the results converge;

[0029] Step 302, using the steady-state flow field as the initial condition of the unsteady-state flow field to perform unsteady-state aerodynamic analysis;

[0030] Step 303, deriving the pressure value of the thermal barrier coating surface of the turbine blade after calculation, and deriving the transient pressure of the point to be analyzed, to obtain the pressure field of the turbine blade surface that changes with the time period.

[0031] In one embodiment, step 302 uses a single-channel method to perform unsteady-state aerodynamic analysis, the solution type is set to a transient rotor row, and a time transformation method is used to reduce the error caused by the phase difference between the dynamic and static fluid domains of a single channel, and the aerodynamic cycle is set to a spacing of the turbine moving blade passing through the turbine static blade, and a number of time steps are set in one aerodynamic cycle.

[0032] In one embodiment, the step 4 is implemented as follows:

[0033] Step 401, performing transient response analysis on the resonance characteristics and critical speed, with the initial and end time and time step settings being consistent with those of the unsteady-state aerodynamic analysis;

[0034] Step 402, interpolating the pressure field that changes with time period to the surface of the thermal barrier coating in the geometric model of the turbine blade containing the thermal barrier coating, importing the steady-state temperature field into the thermal barrier coating and the turbine blade, setting the blade root fixed and rotating speeds, and performing transient response analysis;

[0035] Step 403, deriving the Mises stress, maximum principal stress and displacement values ​​of the surface of the turbine blade geometric model containing the thermal barrier coating and the interface between the blade substrate and the thermal barrier coating over time, that is, the stress field and displacement field at different times.

[0036] In one embodiment, the step 5 specifically includes:

[0037] Step 501, write a fast Fourier transform program;

[0038] Step 502, perform spectral analysis on the pressure, displacement and stress of the point to be analyzed, obtain the amplitude and frequency of the pressure, displacement and stress of the thermal barrier coating at the point to be analyzed under airflow excitation, screen the dangerous position points, obtain the dynamic response amplitude and frequency of the thermal barrier coating at the dangerous position points under airflow excitation, and provide guidance for failure prediction and design of the coating.

[0039] Compared with the prior art, the present invention has the following beneficial effects:

[0040] The present invention realizes the numerical simulation of the vibration stress of the thermal barrier coating of the turbine blade under the upstream stator wake excitation; the aerodynamic load on the surface of the thermal barrier coating is obtained through unsteady calculation, which improves the calculation accuracy; the unsteady flow field adopts the single-channel time transformation method to improve the calculation speed while ensuring the calculation accuracy; the fast Fourier transform method is used to convert the time oscillation of stress into the frequency domain, which is more conducive to analysis.

[0041] In summary, the present invention provides a numerical simulation method for airflow excitation force of thermal barrier coatings, which greatly reduces the cost of studying the damage mechanism of thermal barrier coatings and optimizing the design of thermal barrier coatings, and has good economic benefits. BRIEF DESCRIPTION OF THE DRAWINGS

[0042] Figure 1 It is a flow chart of the numerical simulation analysis method of airflow excitation of thermal barrier coating of turbine blades under high temperature of the present invention.

[0043] Figure 2 The geometric model of the turbine blade with thermal barrier coating established for the present invention, the left picture shows the stationary blade, the right picture shows the moving blade, and the arrow points to the thermal barrier coating.

[0044] Figure 3 The geometric model of turbine stator flow field and the geometric model of turbine moving blade flow field established for the present invention, the left figure is the stator flow field, the right figure is the moving blade flow field, the left arrow represents the grid inlet, the right arrow represents the grid outlet, and the upper and lower sides represent the periodic boundaries.

[0045] Figure 4 It is a Campbell diagram of the modal analysis of the thermal barrier coating of the turbine blade of the present invention, in which a1, a2, a3, a4, a5, a6, and a7 represent the first 7 modal curves respectively.

[0046] Figure 5 for a Point pressure time domain diagram.

[0047] Figure 6 for a Frequency domain plot of point pressure.

[0048] Figure 7 is the equilibrium value of the Mises stress at the interface between the substrate and the thermal barrier coating.

[0049] Figure 8 is the equilibrium value of the maximum shear stress at the interface between the substrate and the thermal barrier coating.

[0050] Fig. 9 for a Point Mises stress time domain diagram.

[0051] Fig.10 for a Point Mises stress frequency domain diagram.

[0052] Fig.11 for a Time domain diagram of the maximum shear stress at the point.

[0053] Fig.12 for a Frequency domain plot of the maximum shear stress at the point. DETAILED DESCRIPTION

[0054] In order to make the purpose, technical scheme and advantages of the present invention clearer, the present invention is further described in detail below in conjunction with specific embodiments and with reference to the accompanying drawings. It should be understood that these descriptions are only exemplary and are not intended to limit the scope of the present invention. In addition, in the following description, the description of well-known structures and technologies is omitted to avoid unnecessary confusion of the concept of the present invention.

[0055] like Figure 1 As shown, the numerical analysis method of airflow excitation of the thermal barrier coating of turbine blades under high temperature of the present invention mainly comprises the following steps:

[0056] Step 1: In the finite element modeling software, establish the external flow field geometry model and the turbine blade geometry model containing the thermal barrier coating.

[0057] Turbine blades include turbine stationary blades (also known as guide blades) and turbine moving blades. In the finite element modeling software, the turbine stationary blades and turbine moving blades are geometrically modeled separately to obtain the turbine stationary blade geometric model and turbine moving blade geometric model, which are recorded as Blade and Vane respectively and saved in .x_t format.

[0058] A thermal barrier coating is established on the turbine blade substrate of the obtained turbine blade geometric model to obtain a turbine blade geometric model containing a thermal barrier coating, which is recorded as TBC and saved in .x_t format. In this embodiment, there is a 0.3 mm thermal barrier coating on the turbine blade, such as Figure 2 shown.

[0059] The geometric model of the turbine stationary blade and the geometric model of the turbine moving blade with the thermal barrier coating constitute the geometric model of the turbine blade with the thermal barrier coating.

[0060] The external flow field geometry model of the present invention is a periodic single-channel fluid geometry model. In this step, a turbine static blade flow field geometry model and a turbine moving blade flow field geometry model are established respectively, which are recorded as Static and Rotor respectively and saved in .x_t format. The geometric positions of the two flow field geometry models are connected to form the external flow field geometry model. The external flow field of the present invention refers to a unit channel of a periodic turbine blade. In this embodiment, the external flow field is specifically a unit channel of a 36:42 periodic turbine blade. The geometry model is as follows: Figure 3 .

[0061] Step 2: Set up the computational grid for each model.

[0062] In this step, the solid domain computational grid and the periodic fluid domain computational grid of the turbine blade thermal barrier coating are established in the grid division software. The specific implementation method is as follows:

[0063] For the solid domain computational mesh of the thermal barrier coating of the turbine blade, the geometric model of the turbine stator blade and the geometric model of the turbine moving blade with the thermal barrier coating are divided into computational meshes. Since the thickness of the geometric model of the turbine moving blade with the thermal barrier coating is much smaller than that of the geometric model of the turbine stator blade, in order to improve the mesh quality, the mesh of the geometric model of the turbine moving blade with the thermal barrier coating needs to be refined. Specifically, the mesh division standard can be set to a mesh quality greater than 0.3. The mesh parts of the geometric model of the turbine moving blade with the thermal barrier coating are named the blade root part Hub, the blade tip part Shroud and the coating part TC, and the interface between the blade base and the thermal barrier coating is named Sub-TC. The Hub, Shroud, TC and Sub-TC parts are all exported to the mesh format of the finite element software used.

[0064] For the periodic fluid domain computational grid, the turbine stator flow field geometric model and the turbine moving blade flow field geometric model are divided into computational grids, and the boundary layer grid is divided on the surface of the blade (including the moving blade and the stator blade) to optimize the quality of the grid. In this embodiment, the standard for dividing the boundary layer grid on the blade surface is Y + <10, where Y + It is a dimensionless quantity that changes according to different flow states. Finally, the grid inlet, outlet, blade surface, and periodic interface are named respectively and exported to the grid format of the fluid calculation software. This embodiment uses the fluid calculation software CFX.

[0065] Step 3: Use finite element software to perform modal analysis on the geometric model of the turbine blade containing the thermal barrier coating to obtain its resonance characteristics and critical speed.

[0066] The specific process of this step can be further described as follows:

[0067] Step 301, import the solid domain calculation mesh of the turbine blade thermal barrier coating obtained in step 2 into the Ansys finite element software, check the mesh, and define the material parameters of the substrate and the thermal barrier coating, including density, elastic modulus, Poisson's ratio, thermal conductivity and specific heat capacity. The parameters used in this embodiment are shown in Table 1.

[0068] Table 1 Material parameters

[0069]

[0070] Step 302, defining the boundary conditions of the geometric model of the turbine blade with the thermal barrier coating, defining fixed boundary conditions, elastic support boundary conditions and contact boundary constraint conditions for the root portion of the turbine blade, and defining other portions of the turbine as free surfaces.

[0071] Among them, the fixed boundary condition is also called rigid constraint, which means that at the fixed point, the displacement (translation and rotation) in all directions is zero; the elastic support boundary condition means that a part of the structure is constrained by elastic support, and the displacement of the structure at the support point is proportional to the reaction force; the contact boundary constraint condition is the interaction between two or more parts of the structure. In the contact area, the parts of the structure can separate or slide but cannot penetrate each other.

[0072] Step 303, after setting the number of modes to be solved and the speed, the natural frequency of the thermal barrier coating of the turbine blade at different speeds is calculated, that is, the Campbell diagram. Through the Campbell diagram, the intersection of the natural frequency and the excitation frequency of the system at different speeds can be determined, so as to find the critical speed ω where resonance occurs, and find the coordinate value of the point to be analyzed from the relative stress and displacement of the thermal barrier coating under resonance. The point to be analyzed can be any point, and the point with larger stress value and displacement can be selected. Dangerous position points can be screened out by setting further thresholds.

[0073] In this embodiment, after setting the first 6 modes and the rotation speed to be solved to 0-20000 RPM, the natural frequency of the thermal barrier coating of the turbine blade at different rotation speeds is calculated, that is, the Campbell diagram, as shown in Figure 4 It can be found that when the rotation speed ω = 718 rad / s, the system may experience the first-order modal bending resonance with a resonance frequency of 4190 Hz. Although resonance may also occur at other rotation speeds, only the case of a rotation speed of 718 rad / s is analyzed here for illustration.

[0074] Step 4: Use fluid analysis software to perform steady-state aerodynamic analysis on the external flow field geometry model to obtain its steady-state flow field and steady-state temperature field. Use the obtained steady-state flow field as the initial condition of the unsteady-state flow field to perform unsteady-state aerodynamic analysis on the external flow field geometry model to obtain the pressure field on the turbine blade surface that changes periodically with time.

[0075] When performing aerodynamic analysis of the external flow field in this step, the steady-state flow field is first solved to obtain the steady-state flow field and steady-state temperature field data of the external flow field. The steady-state flow field is used as the initial condition of the unsteady flow field. Then, the pressure field on the surface of the turbine blade that changes with the time period is calculated. The implementation method is as follows:

[0076] Step 401, import the turbine stator flow field geometry model and turbine moving blade flow field geometry model of the divided calculation grid into the fluid analysis software, define the material of the fluid calculation domain as ideal air, use the Shear Stress Transport turbulence model and the non-equilibrium near-wall model, set the rotation speed to the critical speed ω solved in step 3, set the inlet static pressure, temperature, turbulence and outlet static pressure, and use the frozen rotor method for calculation. The inlet static pressure, temperature, turbulence and outlet static pressure set in this embodiment are as shown in Table 2, set 1000 iteration steps for solution, and obtain the steady-state flow field and steady-state temperature field, i.e., the steady-state result, after the result converges to 0.00001.

[0077] Table 2

[0078] Total inlet temperature 306K Total inlet pressure 169kPa Turbulence 5% Outlet static pressure 158kPa

[0079] Step 402, taking the steady-state flow field as the initial condition of the unsteady-state flow field, using the single-channel method to reduce the amount of calculation, setting the solution type to the transient rotor row, and using the time transformation method to reduce the error caused by the phase difference between the dynamic and static flow domains of the single channel to improve the calculation accuracy, taking the interval of the turbine moving blade passing through the turbine static blade as the aerodynamic cycle, and setting 40 time steps in one cycle.

[0080] Step 403, derive the pressure value of the thermal barrier coating surface of the turbine blade after calculation, and derive the transient pressure of the point to be analyzed, so as to obtain the pressure field of the turbine blade surface that changes with the time period.

[0081] In this embodiment, the point to be analyzed a The pressure at a point changes with time and its spectrum is as follows Figure 5 and Figure 6 It can be found that the vibration of pressure can be decomposed into vibrations of frequencies 4200Hz and 8920Hz, and their amplitudes are 12000Pa and 1362Pa respectively.

[0082] Step 5, interpolating the steady-state temperature field and the pressure field that changes periodically with time to the surface of the thermal barrier coating in the geometric model of the turbine blade containing the thermal barrier coating, performing transient response analysis, and obtaining stress fields and displacement fields at different times.

[0083] The specific implementation method of this step can be described as follows:

[0084] Step 501 , further performing transient response analysis on the resonance characteristics and critical speed of step 3 , with the initial and end time and time step settings being consistent with the flow field transient analysis in step 3 .

[0085] Step 502, interpolate the time-varying pressure field solved in step 4 to the surface of the thermal barrier coating in the geometric model of the turbine blade containing the thermal barrier coating, import the steady-state temperature field into the thermal barrier coating and the turbine blade, set the blade root fixed and rotating speeds, and solve the surface of the geometric model of the turbine blade containing the thermal barrier coating, as well as the Mises stress, maximum principal stress and displacement of the interface between the blade substrate and the thermal barrier coating that vary with time.

[0086] Step 503, export the solution result of step 502 and save it in EXCEL format.

[0087] like Figure 7 and Figure 8 They are respectively the steady-state Mises stress and maximum shear stress of the interface between the blade substrate and the thermal barrier coating of this embodiment. It can be found that the maximum Mises stress is at the leading edge of the blade root, with a maximum value of 102.5 MPa, and the maximum shear stress is at the leading edge of the blade root, with a maximum value of 55.5 MPa.

[0088] Step 6, using fast Fourier transform, the pressure field, stress field and displacement field data of the dangerous position point are transformed into time spectrum to obtain the dynamic response amplitude and frequency of the thermal barrier coating under airflow excitation. The dangerous position point of the present invention refers to the position where the stress is greater than the set value, generally the leading edge area of ​​the blade.

[0089] This step can be specifically described as follows:

[0090] Step 501, use Python, MATLAB and other tools to write a fast Fourier transform program.

[0091] Step 502, through the fast Fourier transform program, the pressure, displacement and stress of the point to be analyzed are subjected to spectral analysis (for analyzing the distribution of the signal on different frequency components), and the amplitude and frequency of the pressure, displacement and stress at the dangerous position of the thermal barrier coating under airflow excitation, that is, the dynamic response amplitude and frequency, are obtained to provide guidance for failure prediction and design of the coating.

[0092] like Fig. 9 , Fig.10 , Fig.11 and Fig.12 They are aThe changes of the point Mises stress and the maximum shear stress with time and their frequency spectrum all show that there is a peak at a frequency of about 4200, and its vibration frequency is close to the airflow vibration frequency, proving that the change of stress is a response to airflow excitation. The amplitude of its vibration is 1.4MPa and 2.5MPa, which is mainly because this embodiment only solves the airflow excitation response of the thermal barrier coating of the turbine blade at low temperature, low pressure and low speed, while the actual turbine engine working conditions are far more complicated than this. The use of this method can solve the airflow excitation response of the thermal barrier coating in a more complex environment, providing important guidance for engineering and design.

Claims

1. A numerical analysis method for airflow excitation of thermal barrier coating of turbine blades under high temperature, characterized in that: The following steps are involved: Step 1, establishing an external flow field geometry model and a turbine blade geometry model with a thermal barrier coating, and setting calculation grids respectively; Step 2, using finite element software to perform modal analysis on the geometric model of the turbine blade containing the thermal barrier coating to obtain its resonance characteristics and critical speed; Step 3, using fluid analysis software to perform steady-state aerodynamic analysis on the external flow field geometric model to obtain its steady-state flow field and steady-state temperature field; using the obtained steady-state flow field as the initial condition of the unsteady-state flow field, using the single-channel method to perform unsteady-state aerodynamic analysis on the external flow field geometric model, the solution type is set to transient rotor row, and the time transformation method is used to reduce the error caused by the phase difference between the dynamic and static fluid domains of the single channel, the aerodynamic cycle is set to a spacing of the turbine moving blade passing through the turbine static blade, and a number of time steps are set in an aerodynamic cycle to obtain the pressure field on the turbine blade surface that changes with the time period; Step 4: The implementation method is as follows: Step 401, performing transient response analysis on the resonance characteristics and critical speed, with the initial and end time and time step settings being consistent with those of the unsteady-state aerodynamic analysis; Step 402, interpolating the pressure field that changes with time period to the surface of the thermal barrier coating in the geometric model of the turbine blade containing the thermal barrier coating, importing the steady-state temperature field into the thermal barrier coating and the turbine blade, setting the blade root fixed and rotating speeds, and performing transient response analysis; Step 403, deriving the Mises stress, maximum principal stress and displacement values ​​of the surface of the turbine blade geometric model with the thermal barrier coating and the interface between the blade substrate and the thermal barrier coating over time, as stress fields and displacement fields at different times; Step 5, using fast Fourier transform, the pressure field, stress field and displacement field data of the dangerous position point are transformed into time spectrum to obtain the dynamic response amplitude and frequency of the thermal barrier coating under airflow excitation. The dangerous position point refers to the position where the stress is greater than the set value.

2. The method for numerical analysis of airflow excitation of thermal barrier coating of turbine blades at high temperature according to claim 1, characterized in that: The step 1 is to establish an external flow field geometry model and a turbine blade geometry model containing a thermal barrier coating in a finite element modeling software, wherein the external flow field geometry model is a periodic single-channel fluid geometry model; and to establish a solid domain computational grid and a periodic fluid domain computational grid of the turbine blade thermal barrier coating in a meshing software.

3. The numerical analysis method for airflow excitation of thermal barrier coating of turbine blades under high temperature according to claim 2 is characterized in that: The external flow field geometric model and the turbine blade geometric model with thermal barrier coating are established, and the implementation method is as follows: Step 101, respectively establishing a turbine stationary blade geometric model and a turbine moving blade geometric model; Step 102, establishing a thermal barrier coating on the turbine blade substrate of the turbine blade geometric model to obtain a turbine blade geometric model containing the thermal barrier coating; the turbine stator blade geometric model and the turbine blade geometric model containing the thermal barrier coating constitute the turbine blade geometric model containing the thermal barrier coating; Step 103, respectively establish a turbine stator flow field geometry model and a turbine rotor flow field geometry model, and connect the geometric positions of the two flow field geometry models to form the external flow field geometry model, and the external flow field is a unit channel of a periodic turbine blade.

4. The method for numerical analysis of airflow excitation of thermal barrier coating of turbine blades at high temperature according to claim 3 is characterized in that: The step 1 is to establish a solid domain computational grid for the thermal barrier coating of the turbine blade, and the implementation method is as follows: Dividing the geometric model of the turbine stator blade and the geometric model of the turbine moving blade with the thermal barrier coating into meshes, naming the mesh parts of the geometric model of the turbine moving blade with the thermal barrier coating as the blade root part, the blade tip part and the coating part, and exporting the blade root part, the blade tip part and the coating part as well as the interface between the blade base and the thermal barrier coating into the mesh format of the finite element software used; The periodic fluid domain computational grid is established as follows: The turbine stator flow field geometry model and the turbine moving blade flow field geometry model are meshed, and the boundary layer mesh is divided on the blade surface. The mesh inlet, outlet, blade surface, and periodic interface are named respectively and exported to the mesh format of the fluid calculation software.

5. The method for numerical analysis of airflow excitation of thermal barrier coating of turbine blades at high temperature according to claim 4, characterized in that: The implementation method of step 2 is as follows: Step 201, importing the solid domain calculation grid of the turbine blade thermal barrier coating into finite element software, and defining the material parameters of the blade substrate and the thermal barrier coating respectively, including density, elastic modulus, Poisson's ratio, thermal conductivity and specific heat capacity; Step 202, defining boundary conditions of a geometric model of a turbine blade with a thermal barrier coating, defining fixed boundary conditions, elastic support boundary conditions and contact boundary constraint conditions for the root portion of the turbine blade, and defining free surfaces for other portions of the turbine blade; Step 203, after setting the number of modes to be solved and the rotation speed, the natural frequency of the thermal barrier coating of the turbine blade at different rotation speeds is calculated, the critical rotation speed ω at which resonance occurs is determined, and the coordinate value of the point to be analyzed is found from the relative stress and displacement of the thermal barrier coating under resonance.

6. The method for numerical analysis of airflow excitation of thermal barrier coating of turbine blades at high temperature according to claim 5, characterized in that: The fixed boundary condition means that at a fixed point, the displacement in all directions is zero; the elastic support boundary condition means that a part of the structure is constrained by elastic support, and the displacement of the structure at the support point is proportional to the reaction force received; the contact boundary constraint condition means the interaction between two or more parts of the structure. In the contact area, the parts of the structure can separate or slide but cannot penetrate each other.

7. The method for numerical analysis of airflow excitation of thermal barrier coating of turbine blades at high temperature according to claim 6, characterized in that: The implementation method of step 3 is as follows: Step 301, import the turbine stator flow field geometry model and turbine moving blade flow field geometry model of the divided calculation grid into the fluid analysis software, define the material of the fluid calculation domain as ideal air, adopt the turbulence model and the non-equilibrium near-wall model, set the inlet static pressure, temperature, turbulence and outlet static pressure, adopt the frozen rotor method for calculation, set the iterative step solution, and obtain the steady-state flow field and steady-state temperature field after the results converge; Step 302, using the steady-state flow field as the initial condition of the unsteady-state flow field to perform unsteady-state aerodynamic analysis; Step 303, deriving the pressure value of the thermal barrier coating surface of the turbine blade after calculation, and deriving the transient pressure of the point to be analyzed, to obtain the pressure field of the turbine blade surface that changes with the time period.

8. The method for numerical analysis of airflow excitation of thermal barrier coating of turbine blades at high temperature according to claim 1, characterized in that: The step 5 specifically includes: Step 501, write a fast Fourier transform program; Step 502, perform spectral analysis on the pressure, displacement and stress of the point to be analyzed, obtain the amplitude and frequency of the pressure, displacement and stress of the thermal barrier coating at the point to be analyzed under airflow excitation, screen the dangerous position points, obtain the dynamic response amplitude and frequency of the dangerous position points of the thermal barrier coating under airflow excitation, and provide guidance for failure prediction and design of the coating.

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