A fatigue crack growth analysis method for rocket engine turbine blades

Abstract

The present invention discloses a fatigue crack propagation analysis method for rocket engine turbine blades, including establishing a rocket engine turbine blade fatigue crack propagation model; determining high and low cycle fatigue loads and inputs through operating loads; combining ABAQUS and FRANC3D to calculate the fatigue crack propagation of turbine blades; outputting fatigue fracture parameters such as stress intensity factor and J integral to determine the critical crack size of turbine blades; and predicting the high and low cycle fatigue life of the rocket engine turbine blades according to the fatigue crack propagation model. Compared with the existing fatigue crack propagation analysis, the present invention takes into account the operating loads of the rocket engine turbine blades during the start-stop, stable operation, and overload full cycle process, realizes the calculation of the high and low cycle composite fatigue problem of the rocket engine turbine blades, and can more accurately predict the fatigue crack propagation life and critical crack length of the rocket engine turbine blades under real working conditions.

Description

Technical Field

[0001] The invention relates to blade fatigue crack growth prediction, in particular to a method for analyzing fatigue crack growth of a rocket engine turbine blade. Background Art

[0002] As a key component of a rocket engine, turbine blades are required to work safely and reliably in a high-temperature, high-load, and high-speed working environment. During operation, the blades are impacted by periodic high-speed airflow excitation forces, and their fatigue problems are very prominent and complex. Due to the short working time of existing liquid rocket engine turbines, the design process mainly considers aerodynamic performance and static strength, and rarely considers blade vibration fatigue and high and low cycle fatigue problems caused by gas excitation. However, with the development of reusable engines, the service life of various components is required to increase exponentially. Therefore, it is necessary to conduct in-depth research on the high and low cycle fatigue problems of turbine blades. The core of this problem not only involves the analysis of high and low cycle composite fatigue problems of blades under huge centrifugal static loads and airflow excitation forces, but also involves the determination of the critical crack size of high-temperature rotating components.

[0003] Fracture mechanics can predict and estimate the safety of materials and structures, and is the development direction of fatigue performance analysis of aerospace engine turbine blades. Related simulations can evaluate the fatigue life of structures with existing cracks, showing the advanced nature of focusing on mechanism analysis and life prediction, and have been widely used in aviation, aerospace, machinery and other structures. Tools and methods based on fracture mechanics can more reasonably and accurately predict fatigue crack growth life and crack growth path. FRANC3D crack analysis software mainly calculates the crack growth and fatigue life of three-dimensional cracks. At present, this software has been used to carry out a lot of crack growth analysis work, proving that it is a reliable means of crack growth and life prediction. In the prior art, a method for estimating the remaining fatigue life of aluminum-lithium alloy repair structures in the aerospace field is proposed based on ABAQUS combined with FRANC3D. There is still a lack of crack growth analysis of key positions of liquid rocket engine turbine blades under full-cycle service environment under high and low cycle combined fatigue.

[0004] In terms of blade crack propagation theory and numerical analysis, current research is mostly limited to surface cracks and through-cracks. However, in reality, blades mostly exist in the form of three-dimensional cracks. The modeling, numerical simulation, and theoretical analysis of the crack front morphology, crack propagation path, and crack propagation life are very complex. The current analysis of blade crack propagation is still not accurate enough. Summary of the invention

[0005] Purpose of the invention: In view of the above problems, the present invention provides a fatigue crack propagation analysis method for rocket engine turbine blades with high simulation accuracy.

[0006] Technical solution: To solve the above problems, the present invention adopts a fatigue crack growth analysis method for a rocket engine turbine blade, comprising the following steps:

[0007] Step 1: Perform finite element analysis pre-processing in ABAQUS software to establish a finite element model of a local intake impulse turbine blade, including establishing material property parameters, mesh refinement, load conditions and interactions that meet working conditions;

[0008] Step 2: Import the actual working condition of the local air intake impulse turbine during normal operation into the load condition, establish an analysis step including all stages of operation of the local air intake impulse turbine blade, perform load setting in the analysis step of each stage, perform static analysis on the finite element model of the local air intake impulse turbine blade, and obtain the location of the danger point;

[0009] Step 3: In FRANC3D software, the local intake impact turbine blade finite element model set in step 2 is divided into a local crack extension model and the remaining part, an initial crack is set at the dangerous point of the local crack extension model, a high-low cycle composite fatigue load is set for the local crack extension model using non-proportional loading, and a high-low cycle composite fatigue crack extension calculation is performed;

[0010] Step 4: Calculate the local crack propagation model set in step 3 in the FRANC3D software, output the stress intensity factor and J-integral fracture parameter, calculate the critical crack length of the local intake impulse turbine blade according to the residual strength theory, and calculate the final fatigue crack propagation life of the local intake impulse turbine blade through the fatigue crack growth rate model.

[0011] In the above, the specific method of step 1 is:

[0012] A local intake impulse turbine blade model is established using 3D modeling software and imported into ABAQUS software to establish material property parameters, mesh refinement, load conditions and interactions that meet working conditions, and obtain a finite element model of the local intake impulse turbine blade.

[0013] Furthermore, the finite element model of the local inlet impingement turbine blade was meshed using ABAQUS software, using C3D10 second-order elements, which were defined by 10 nodes, each with three degrees of freedom. Mesh encryption was performed at the dangerous point (where the blade has a variable cross section).

[0014] In the above, the specific method of step 2 is:

[0015] The local air intake impulse turbine blades are subjected to the centrifugal force brought by the rotation when working. Therefore, the loading settings of the finite element model of the local air intake impulse turbine blades in ABAQUS software include: setting the rotating body force and importing the speed information of each stage; setting the temperature predefined field and importing the temperature information of each stage; setting the pressure loading and mapping the given aerodynamic load data to the finite element node information to realize the setting of the temperature, speed, aerodynamic load and boundary conditions of the local air intake impulse turbine blades at each stage.

[0016] Specifically, in ABAQUS software, the blade temperature field and aerodynamic load are set as predefined fields using mapping fields, acting on the global model of a single blade, and enabled in the analysis step Load_Step3. At the same time, the analysis step Load_Step1 is set to the initial state, the speed is zero, and there is no aerodynamic and temperature load; the analysis step Load_Step2 is set to centrifugal static load, the speed is 19331rpm, the temperature is the temperature field under the actual working conditions, and there is no aerodynamic load; Load_Step3 is set to centrifugal static load, and the temperature field and aerodynamic load under the actual working conditions are applied at the same time.

[0017] After importing the real working conditions of the local intake impulse turbine during normal operation, the static force analysis of the finite element model of the local intake impulse turbine blade was carried out in ABAQUS software, and the specific location of stress concentration was found to be the variable cross-section at the root of the blade, and the location with the maximum stress was taken as the danger point.

[0018] In the above, the specific method of step 3 is:

[0019] The inp file of the local intake impulse turbine blade finite element model established by ABAQUS is imported into FRANC3D software to divide the local crack propagation model. The initial crack is set at the dangerous point, and the initial crack morphology and type are used as influencing factors of fatigue life and crack propagation.

[0020] Furthermore, a local air intake impulse turbine will go through three processes during one use: startup, stable operation, and shutdown. Startup and shutdown involve strong transients and thermal shocks from zero to tens of thousands of revolutions per second, which constitute typical low-cycle fatigue loads; while during stable operation, it is subject to the impact of gas, and the alternating aerodynamic load constitutes a typical high-cycle fatigue load.

[0021] In FRANC3D, non-proportional loading is used to set high- and low-cycle combined fatigue loads. When calculating high-cycle fatigue, the aerodynamic load is superimposed on the centrifugal static load in the stable operation stage as the maximum stress intensity factor K. max , the centrifugal static load in the stable operation stage is taken as the minimum stress intensity factor K minThe maximum stress intensity factor corresponds to the load state of maximum aerodynamic pressure plus centrifugal force, the minimum stress intensity factor corresponds to the load state of pure centrifugal force, and the stress intensity factor amplitude is the difference between the two. Low-cycle fatigue corresponds to the process from startup to overspeed, plus the two stages from overspeed to stable operation.

[0022] Furthermore, when crack propagation analysis is performed on the local crack propagation model of high-temperature rotating components in FRANC3D, the maximum tensile stress criterion is selected for crack propagation torsion angle calculation, crack propagation is simulated by a given crack front extension amount, and the parameters use the temperature-dependent NASGRO model and parameters. When the minimum stress intensity factor of the crack front reaches the threshold value, the crack begins to propagate. When the crack stress intensity factor reaches the fracture toughness of the material, or when the calculation model cannot proceed, the calculation stops;

[0023] The NASGRO model takes into account the three stages of crack extension and the crack closure effect, and its crack extension formula is:

[0024] ;

[0025] in, is the number of fatigue loading cycles, is the crack length, is the stress ratio, , , , is a material-related empirical parameter, is the critical stress intensity factor, is the stress intensity factor threshold, is the stress intensity factor range, is the maximum stress intensity factor, is the crack opening formula, namely:

[0026] ;

[0027] ;

[0028] ;

[0029] ;

[0030] ;

[0031] ;

[0032] ;

[0033] ;

[0034] in, is the ratio of the maximum applied stress to the flow stress, is the plane stress / strain constraint factor, for each specific material, Has been processed into a constant; is the yield limit, is the tensile limit, is the fracture toughness, It's the thickness.

[0035] In the above, the specific method of step 4 is:

[0036] The fracture mechanics parameters of the local crack growth model are calculated, and the stress intensity factor K can be obtained by using the M integral calculation. I , K II , K III The fatigue crack growth life of the whole local intake impulse turbine blade is calculated by NASGRO model. The stress intensity factor criterion and J-integral fracture criterion are used to analyze the crack growth and residual strength of the key parts of the turbine blade to determine the critical crack length and residual strength level.

[0037] The M integral method is the most accurate and commonly used method in FRANC3D software. The present invention uses the M integral for calculation.

[0038] According to different cracking modes, three stress intensity factors can be obtained as follows:

[0039] Type I crack:

[0040] ;

[0041] Type II crack:

[0042] ;

[0043] Type III crack:

[0044] ;

[0045] Where f, g, and h are shape coefficients.

[0046] In fracture mechanics, in order to analyze the stress and strain field around defects / cracks, the line integral J-integral, which is independent of the integral path, is often used. The J-integral has a clear physical meaning: for linear elastic plane crack problems, the J-integral is the strain energy release rate at the crack tip. The purpose of crack growth analysis is to estimate the crack growth life of the damaged structure during the entire safe service life. Under the action of fatigue load, the damage of the object being analyzed will continue to grow, and the crack will grow from the initial crack length to the maximum length. Critical crack length that allows termination The number of load cycles required within this period of time is called the crack growth life.

[0047] The present invention also adopts a computer device, including a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor implements the steps of the above method when executing the computer program.

[0048] The present invention also provides a computer-readable storage medium having a computer program stored thereon, wherein the computer program implements the steps of the above method when executed by a processor.

[0049] Beneficial effects: Compared with the prior art, the significant advantage of the present invention is that it develops a fatigue crack propagation analysis method for rocket engine turbine blades by combining the methods of ABAQUS and FRANC3D. This method can take into account the high- and low-cycle composite fatigue problems of liquid rocket engine turbine blades during startup, operation, and shutdown, accurately simulate the expansion behavior of turbine blade cracks under real working conditions, more accurately predict the service life of the turbine under full-cycle operation, and finally establish a method for determining the critical crack size of high-temperature rotating parts, which helps to clarify the failure mechanism and crack propagation mechanism of high-temperature rotating turbine blades of rocket engines. BRIEF DESCRIPTION OF THE DRAWINGS

[0050] Figure 1 It is a schematic diagram of the process of the present invention.

[0051] Figure 2 It is a schematic diagram of the working principle of the local air intake impulse turbine in the present invention.

[0052] Figure 3 It is a stress response curve diagram at different positions of the root section of the turbine blade in the present invention.

[0053] Figure 4 This is a morphology diagram of the initial crack and fatigue crack extension at the back of the turbine blade in the present invention.

[0054] Figure 5 is the stress intensity factor K of the crack front edge at the back of the blade in the present invention I Crack extension curve.

[0055] Figure 6 This is a curve diagram showing how the crack length at the blade back and the leading edge of the blade changes with the number of load cycles in the present invention.

[0056] Figure 7 The figure shows the determination of critical crack size based on the stress intensity factor criterion in the present invention.

[0057] Figure 8 This is a diagram showing the determination of critical crack size based on the J-integral criterion in the present invention. DETAILED DESCRIPTION

[0058] like Figure 1 As shown, a fatigue crack growth analysis method for a rocket engine turbine blade in this embodiment comprises the following steps:

[0059] Step 1: Use 3D modeling software to build a local intake impulse turbine blade model, see Figure 2 . The turbine rotor adopts an integrated design of disk, shaft and blades. The blades are straight blades with an aspect ratio of 2. The turbine mid-diameter is 150mm, the blade width is 13mm, and the number of blades is 53. The turbine rotor material is a nickel-based high-temperature alloy. The blank is formed by the integral forging method, and the rotor blades are processed by integral electrolytic machining. A single local intake impact turbine blade model is selected and imported into the ABAQUS software to establish material property parameters, suitable grids, load conditions, interactions, etc. that meet the working conditions. The mechanical property parameters of the blade material change with temperature, as shown in the table below:

[0060]

[0061] The finite element model of the turbine blade uses C3D10 second-order units, which are defined by 10 nodes, each with three degrees of freedom. Mesh encryption is performed at the dangerous point (where the upper and lower sections of the blade change) to ensure the accuracy of the finite element stress and strain calculations. The entire turbine blade consists of 53 blades, so it is necessary to set periodic boundary conditions for each blade, and select cyclic symmetry in the interaction module, where the main contact surface and the secondary contact surface are selected as the two surfaces that contact each other between the blades, and the total number of sectors is set to 53.

[0062] Step 2: Further, analyze the entire working cycle process, including the three stages of startup, stable stage, and shutdown. Consider the high- and low-cycle composite fatigue crack extension and divide the entire process into high-cycle fatigue load and low-cycle fatigue load. According to the rain flow counting method, calculate the stress intensity factor. The single-use high-cycle fatigue load only considers the temperature, speed and aerodynamic load in the stable operation stage, where the speed is a static load. The high- and low-cycle fatigue loads used multiple times consider the entire process, where startup and shutdown are set as low-cycle loads (cycle once), and the load in the stable operation stage is set as high-cycle fatigue, including centrifugal static loads and aerodynamic loads.

[0063] To input the full duty cycle load spectrum in FRANC3D, four analysis steps need to be established in ABAQUS, corresponding to the K of low-cycle fatigue and high-cycle fatigue. min and K max. For the analysis step type, select static general, enable geometric nonlinearity, time length, incremental step and other settings, and design them according to your requirements. Import the real working conditions into ABAQUS, set the rotating body force, and import the speed information of each stage; set the temperature predefined field, and import the temperature information of each stage; set the pressure loading, and map the given aerodynamic load data to the finite element node information (see Figure 3 ). Realize the setting of temperature, speed, aerodynamic load and boundary conditions of turbine blades at each stage.

[0064] In ABAQUS, the temperature field and aerodynamic load of the blade are set as predefined fields by using mapping fields, acting on the global model of a single blade, and set to be enabled in the analysis step Load_Step3. At the same time, the analysis step Load_Step1 is set to the initial state, the speed is zero, and there is no aerodynamic and temperature load; the analysis step Load_Step2 is set to centrifugal static load, the speed is 19331rpm, the temperature is the temperature field under the actual working conditions, and there is no aerodynamic load; Load_Step3 is set to centrifugal static load, and the temperature field and aerodynamic load under the actual working conditions are applied at the same time. After setting the load conditions, static analysis is performed to find out the dangerous nodes and stress-strain conditions. The dangerous nodes are located at the variable cross-sections of the blade back and the leading edge. The maximum stresses of the blade back during the power-on and power-off and stable operation stages are 832 and 359MPa, respectively, and the maximum stresses of the blade leading edge during the power-on and power-off and stable operation stages are 1007 and 543MPa, respectively.

[0065] Step 3: Import the global model inp file of a single turbine blade generated in step 2 into FRANC3D for local crack propagation model division. Set initial cracks at dangerous nodes (where stress is maximum on the blade back and leading edge). The initial crack size is determined according to the GH4169 threshold value. Insert elliptical initial cracks at the maximum stress positions on the blade back and leading edge, respectively. The crack opens in the radial direction. The initial crack sizes at the blade back and leading edge are 0.5 mm and 0.56 mm, respectively.

[0066] In FRANC3D, non-proportional loading is used to set high- and low-cycle combined fatigue loads. When calculating high-cycle fatigue, the aerodynamic load is superimposed on the centrifugal static load in the stable operation stage as K. max , the centrifugal static load in the stable operation stage is taken as K min The maximum stress intensity factor corresponds to the load state of the maximum aerodynamic pressure plus the centrifugal force, the minimum stress intensity factor corresponds to the load state of the pure centrifugal force, and the stress intensity factor amplitude is the difference between the two. Low-cycle fatigue corresponds to the process from startup to overspeed, plus overspeed to stable operation to calculate fatigue crack growth. See the following table for details:

[0067]

[0068] The fatigue crack growth calculation of FRANC3D is heavily dependent on the material fatigue crack growth parameters and models. In the present invention, the temperature-dependent fatigue crack growth parameters of GH4169 in the NASGRO database are selected, as shown in the following table:

[0069]

[0070] The fracture mechanics parameters of the sub-model with cracks are calculated, and the stress intensity factor K can be obtained by using the M integral calculation. I , K II , K III and J integral value. The initial crack morphology and fatigue crack extension morphology are shown in Figure 4 It can be seen that the fatigue crack propagation at the back of the blade is dominated by the opening type (type I) crack, accompanied by the sliding type (type II) and the tearing type (type III) crack. The crack starts to propagate vertically and straightly in the radial direction, and then gradually deflects toward the root. The calculated crack front stress intensity factor K I As the fatigue crack propagation process changes Figure 5 As shown. Stress intensity factor K I The distribution is high at both ends and low in the middle. The stress intensity factor of the crack at the surface is slightly larger. The crack growth rate at both ends is greater than that in the middle. The elliptical crack front gradually changes to a straight crack front. At the same time, the crack growth rate on the left side is slightly greater than that on the right side (intake side). As the number of crack growth steps increases, K I The stress intensity factor increased from the initial 750 Increased to 3400 at the final failure , until the fracture toughness K IC .

[0071] Figure 6 The curves showing the variation of crack length with the number of load cycles at the blade back (point C) and the leading edge (point A) of the blade under the full cycle load are shown. Compared with point A at the leading edge, when the fatigue crack expands from the initial critical crack length to the failure length, the expansion speed at point C at the blade back is faster, the crack expansion length is shorter, and the fatigue crack expansion life is shorter. Under this condition, the crack expansion life of the blade back of the local intake impulse turbine is 20 times, which is 5 less than that at the leading edge. The reason for this is analyzed as follows: the ΔKI value of the blade back under high and low cycle fatigue loads is greater than that of the leading edge, and the stress ratio is also greater than that of the leading edge, which makes the crack expansion life at the blade back calculated based on the NASGRO fatigue crack expansion model shorter than that at the leading edge.

[0072] Based on the residual strength theory, a uniform 15% residual strength margin is set, that is, it is considered that the structural component fails when it reaches 85% of the maximum strength, and the corresponding crack length is the critical crack length. Figure 7 and Figure 8 The fracture toughness (K IThe critical crack lengths at the back of the blade determined by the stress intensity factor K and J integral are 2.21 mm and 2.29 mm respectively. I The critical crack length result is too large.

Claims

1. A method for analyzing fatigue crack growth of a rocket engine turbine blade, characterized in that: The following steps are involved: Step 1: Perform finite element analysis pre-processing in ABAQUS software to establish a finite element model of a local intake impulse turbine blade, including establishing material property parameters, mesh refinement, load conditions and interactions that meet working conditions; Step 2: Import the actual working condition of the local air intake impulse turbine during normal operation into the load condition, establish an analysis step including all stages of operation of the local air intake impulse turbine blade, perform load setting in the analysis step of each stage, perform static analysis on the finite element model of the local air intake impulse turbine blade, and obtain the location of the danger point; Step 3: In FRANC3D software, the local intake impact turbine blade finite element model set in step 2 is divided into a local crack extension model and the remaining part, an initial crack is set at the dangerous point of the local crack extension model, a high-low cycle composite fatigue load is set for the local crack extension model using non-proportional loading, and a high-low cycle composite fatigue crack extension calculation is performed; Step 4: Calculate the local crack growth model set in step 3 in FRANC3D software, output the stress intensity factor and J-integral fracture parameter, calculate the critical crack length of the local air intake impulse turbine blade according to the residual strength theory, and calculate the final fatigue crack growth life of the local air intake impulse turbine blade through the fatigue crack growth rate model; The specific method of step 3 is as follows: The inp file of the local air intake impulse turbine blade finite element model established by ABAQUS software is imported into FRANC3D software to divide the local crack extension model, and the initial crack is set at the dangerous point. The morphology and type of the initial crack are used as the influencing factors of fatigue life and crack extension. In FRANC3D software, non-proportional loading is used to set high- and low-cycle composite fatigue loads to simulate the working conditions of local air intake impulse turbine blades, including high-cycle fatigue loads during stable operation and low-cycle fatigue loads during startup and shutdown of the local air intake impulse turbine. When calculating the high-cycle fatigue loads on the local air intake impulse turbine blades, the aerodynamic load of the local air intake impulse turbine and the centrifugal static load in the stable operation stage are superimposed as the maximum stress intensity factor K. max , the centrifugal static load of the local intake impulse turbine during the stable operation phase is taken as the minimum stress intensity factor K min The maximum stress intensity factor corresponds to the load state of the maximum aerodynamic pressure plus the centrifugal force, the minimum stress intensity factor corresponds to the load state of the pure centrifugal force, the stress intensity factor amplitude is the difference between the maximum stress intensity factor and the minimum stress intensity factor, and the low-cycle fatigue load includes the process from startup to overspeed, plus the two stages of overspeed to stable operation; When performing fatigue crack growth calculations for the local crack growth model, the crack growth is simulated using a given crack front growth amount, using the temperature-dependent NASGRO model and parameters. When the minimum stress intensity factor at the crack front reaches the threshold value, the crack begins to grow. When the crack stress intensity factor reaches the fracture toughness of the material, or when the calculation model cannot proceed, the calculation stops. The NASGRO model takes into account the three stages of crack growth and the crack closure effect. The crack growth formula is: ； in, is the number of fatigue loading cycles, is the crack length, is the stress ratio, , , , is a material-related empirical parameter, is the critical stress intensity factor, is the stress intensity factor threshold, is the crack opening formula, is the stress intensity factor range, is the maximum stress intensity factor.

2. The fatigue crack growth analysis method of a rocket engine turbine blade according to claim 1, characterized in that: The specific method of step 1 is as follows: A local air intake impulse turbine blade model is established through 3D modeling software and imported into ABAQUS software. The material property parameters, mesh refinement, load conditions and interactions that meet the working conditions are established to obtain the finite element model of the local air intake impulse turbine blade. The finite element model of the local air intake impulse turbine blade is meshed using ABAQUS software. The C3D10 second-order unit is used. The unit is defined by 10 nodes, each node has three degrees of freedom, and the mesh is encrypted at the variable section of the finite element model of the local air intake impulse turbine blade.

3. The fatigue crack growth analysis method of a rocket engine turbine blade according to claim 2, characterized in that: The specific method of step 2 is as follows: The loading settings for the finite element model of the local intake impulse turbine blade in ABAQUS software include: setting the rotating body force and importing the speed information of each stage; setting the temperature predefined field and importing the temperature information of each stage; setting the pressure loading and mapping the given aerodynamic load data to the finite element node information to realize the setting of the temperature, speed, aerodynamic load and boundary conditions of the local intake impulse turbine blade at each stage.

4. The method for analyzing fatigue crack growth of a rocket engine turbine blade according to claim 3, characterized in that: The specific method of step 2 is as follows: After importing the real working conditions of the local intake impulse turbine during normal operation, the static force analysis of the finite element model of the local intake impulse turbine blade was carried out in ABAQUS software, and the specific location of stress concentration was found to be the variable cross-section at the root of the blade, and the location with the maximum stress was taken as the danger point.

5. The method for analyzing fatigue crack growth of a rocket engine turbine blade according to claim 4, characterized in that: The specific method of step 4 is as follows: The fracture mechanics parameters of the local crack growth model are calculated, and the stress intensity factor K is obtained by using the M integral calculation. I , K II , K III The fatigue crack growth life of the whole local intake impulse turbine blade is calculated by NASGRO model, and the stress intensity factor criterion and J-integral fracture criterion are used to analyze the crack growth and residual strength of the key parts of the local intake impulse turbine blade to determine the critical crack length and residual strength level.

6. A computer device comprising a memory, a processor and a computer program stored in the memory and executable on the processor, characterized in that: When the processor executes the computer program, the steps of the method according to any one of claims 1 to 5 are implemented.

7. A computer-readable storage medium having a computer program stored thereon, characterized in that: When the computer program is executed by a processor, the steps of the method according to any one of claims 1 to 5 are implemented.

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