A method for evaluating the endurance damage of turbine blades based on flight climb mission load spectrum

Abstract

The application relates to the technical field of aero-engines and discloses a turbine blade endurance damage evaluation method based on a flight climb task load spectrum, which is characterized in that the load spectrum in the flight climb stage is divided into regions, the first principal stress, the second principal stress, the third principal stress and the equivalent stress of each node of a turbine blade finite element model are calculated to obtain the equivalent stress of an examination position, the endurance fracture time of the examination position is calculated by using a creep endurance and thermal strength comprehensive parameter unified equation, and finally, the damage is calculated by using Miner linear accumulation, the situation that the engine physical rotating speed, aerodynamic thermal force, internal flow and temperature are reduced due to fuel consumption in the climb stage is considered, the equivalent stress in the creep endurance and thermal strength comprehensive parameter unified equation for endurance fracture time calculation considers the influence of stress relaxation effect, and is optimized by using a multi-target genetic optimization algorithm, the fitting parameters at different temperatures are unified, and the predicted linear cumulative damage result is more accurate.

Description

Technical Field

[0001] This invention relates to the field of aero-engine technology and discloses a method for assessing persistent damage to turbine blades based on the flight climb mission load spectrum. Background Technology

[0002] Flight spectrum is the proportion of statistical data on various flight and ground states and their durations throughout an aircraft's entire lifespan. It comprehensively considers the combination of total aircraft weight, center of gravity position, flight altitude, and flight time, and is a prerequisite for aero-engine design and component fatigue life determination. Based on operational requirements, UAVs need to climb to several specific altitudes during long-endurance flights at medium to high altitudes. Current aerial measurements and statistics on actual UAV flight loads reveal that during prolonged cruise at medium to high altitudes at constant Mach numbers, fuel consumption leads to a reduction in aircraft weight and thrust demand, causing the engine's performance to gradually decline at the same cruise altitude, resulting in a decrease in overall load.

[0003] Currently, engine structural strength design primarily relies on the design load spectrum, derived from experience using the load spectra of in-service aircraft with similar missions and structures. However, due to differences in engine type and purpose, the damage to key engine components, especially turbine blades operating under high stress and high temperature, varies significantly across different flight mission profiles. Based on overall performance parameters from the high-altitude climb phase of typical UAV flight profiles, it was found that engine speed and turbine inlet temperature decrease to some extent during cruise at the same altitude. Designs often fail to consider these variations, typically evaluating turbine blade lifespan based on design load and operating time, neglecting the actual decrease in engine speed and turbine inlet temperature during flight. This results in turbine blade lifespans deviating significantly from design parameters, leading to overly conservative predictions. Summary of the Invention

[0004] The purpose of this invention is to provide a method for assessing persistent damage to turbine blades based on the load spectrum of flight climb missions, which can improve the accuracy of load analysis and stress calculation results, and make the predicted linear cumulative damage results more accurate.

[0005] To achieve the above-mentioned technical effects, the technical solution adopted by the present invention is as follows:

[0006] A method for assessing persistent damage to turbine blades based on the load spectrum of a flight climb mission includes: The load spectrum of the flight climb phase to be analyzed is divided into multiple regions, and the rotational speed and cumulative time of each region are extracted. Each region is divided into multiple load steps. The overall performance data of the engine, air system data, turbine aerodynamic data and material heat transfer performance data under each load step are used as boundary conditions to perform finite element heat transfer analysis of the turbine blades, obtain three-dimensional temperature data of the turbine blades, and obtain S2 aerodynamic data of the turbine blades through aerodynamic analysis. Using the highest physical rotational speed, three-dimensional temperature data of turbine blades, S2 aerodynamic data, and mechanical property data of turbine blade materials in each region as inputs, stress calculation and analysis are performed using a finite element model of turbine blades to obtain the first principal stress, second principal stress, third principal stress, and equivalent stress at the test location of the finite element model in the corresponding load step; based on the first principal stress, second principal stress, third principal stress, and equivalent stress at the test location in the corresponding load step, the stress correction stress at the test location in the corresponding load step is calculated. Using a standard round bar of turbine blade material as the test specimen, the creep rupture time of the test specimen was obtained under different test temperatures and test stresses. Using test temperature and test stress as inputs and creep rupture time as output, a unified equation for the comprehensive parameters of creep rupture and thermal strength of turbine blade material was established. Based on the corrected stress at the test location under the corresponding load step in each region, the equivalent stress at the test location in the corresponding region was obtained by analyzing the unified equation for the comprehensive parameters of creep rupture and thermal strength. Based on the equivalent stress and temperature at the turbine blade test location in each region, the creep fracture time at the test location is obtained by analyzing a unified equation that combines creep and thermal strength parameters. Based on the sustained fracture time at the test location, the linear cumulative damage at the test location of the turbine blade is calculated using the Miner linear cumulative damage model.

[0007] Furthermore, using 1% of the initial physical rotational speed of the flight climb phase to be analyzed as a step, the load spectrum of the flight climb phase to be analyzed is divided into multiple regions.

[0008] Furthermore, methods for obtaining the first principal stress, second principal stress, third principal stress, and equivalent stress at the corresponding load step in the finite element model assessment location include: A creep damage constitutive model for turbine blade materials is constructed, comprising a strain hardening model simulating the first stage of creep damage in turbine blade materials. And the Norton model for simulating the second stage of creep damage in turbine blade materials. ;in To determine the creep strain rate for the first stage of creep damage under the selected test temperature and stress, The test stress selected for the standard round bar of turbine blade material. The test temperature selected for the standard round bar of turbine blade material. To determine the creep strain in the first stage of creep damage under selected test temperature and stress, It is a natural constant. To determine the creep strain rate of the second stage of creep damage under selected test temperature and test stress, to These are the fitting parameters, obtained through data fitting; Based on the tensile test of a standard round bar for turbine blade materials, a true stress-strain curve for the turbine blade material is constructed. This true stress-strain curve, along with the parameters of the fitted creep damage constitutive model, is then analyzed. to Input the finite element model of the turbine blade; Using the highest physical rotational speed of each region, three-dimensional temperature data of turbine blades, S2 aerodynamic data, and mechanical property data of turbine blade materials as inputs, stress analysis is performed using a finite element model of turbine blades to obtain the first principal stress, second principal stress, third principal stress, and equivalent stress at the test location under the corresponding load step.

[0009] Furthermore, the corrected stress at the test location under the corresponding load step. ;in, Indicates the assessment position is at the _____. i The first region j Under each load step t Correction stress at time, , m This represents the total number of regions during the climbing phase. , For the first i Total number of load steps in each region The assessment position is in the 1st place. i The first region j Under each load step t The first principal stress at time t, The assessment position is in the 1st place. i The first region j Under each load step t The second principal stress at time t, The assessment position is in the 1st place. i The first region j Under each load step t The third principal stress at time t, The assessment position is in the 1st place. i The first region j Under each load step t The equivalent force at time, The assessment position is in the 1st place. i The first region j Under each load step t Stress triaxiality at time t, These are empirical coefficients, with values ​​ranging from 0 to 1.

[0010] Furthermore, the assessment position is in the... i The first region j Stress triaxiality under each load step .

[0011] Furthermore, the equivalent stress at the test location in the corresponding region. ,in The assessment position is in the 1st place. i Equivalent stress in each region For the first i Cumulative time in the region For the first i The first in the region End time of load step Correction stress at time, , For the first i Number of load steps in the region For the first i The first in the region j The cumulative time of the load step. The assessment position is in the 1st place. i The power exponent of the region, , To analyze the turbine blade test position obtained by using a unified equation that combines creep and thermal stress parameters. i The first in the region x Corresponding temperature and corrected stress of the load step The duration of sustained fracture, Represents the differential symbol.

[0012] Furthermore, the unified equation for the comprehensive parameters of creep durability and thermal strength of turbine blade materials is established as follows: ,in, The sustained fracture stress of a standard round bar for turbine blade material at a selected test temperature. The test temperature selected for the standard round bar of turbine blade material. The time to sustained fracture of a standard round bar of turbine blade material at a selected test temperature and sustained fracture stress; , , , , , , , These are the model parameters, which are obtained by using a multi-objective genetic optimization algorithm with the goal of minimizing the preset loss function.

[0013] Furthermore, linear cumulative damage at the turbine blade testing location. ,in, For the first i Cumulative time in the region For the first i The turbine blade test location within the region is at the corresponding temperature equivalent stress. The cumulative time below, , m This represents the total number of regions during the climbing phase.

[0014] Compared with the prior art, the beneficial effects of this invention are as follows: This invention takes into account the fact that the engine's physical speed, aerodynamic thermodynamics, internal flow and temperature will decrease due to fuel consumption during the climb phase of an actual flight mission, which improves the accuracy of the analysis load and stress calculation results. The equivalent stress in the equation that unifies the comprehensive parameters of creep and thermal intensity takes into account the influence of stress relaxation effect, and a multi-objective genetic optimization algorithm is used for optimization, which unifies the fitting parameters at different temperatures, making the predicted linear cumulative damage results more accurate. Attached Figure Description

[0015] Figure 1 This is a flowchart of the turbine blade persistent damage assessment method based on the flight climb mission load spectrum in Example 1 or 2. Figure 2 The load spectrum of the flight climb phase to be analyzed in Example 2 A diagram showing the division of the regions. Detailed Implementation

[0016] The present invention will now be described in further detail with reference to the embodiments and accompanying drawings. However, this should not be construed as limiting the scope of the above-described subject matter of the present invention to the following embodiments; all technologies implemented based on the content of the present invention fall within the scope of the present invention.

[0017] Example 1 See Figure 1 A method for assessing persistent damage to turbine blades based on the load spectrum of a flight climb mission, comprising: The load spectrum of the flight climb phase to be analyzed is divided into multiple regions, and the rotational speed and cumulative time of each region are extracted. Each region is divided into multiple load steps. The overall performance data of the engine, air system data, turbine aerodynamic data and material heat transfer performance data under each load step are used as boundary conditions to perform finite element heat transfer analysis of the turbine blades, obtain three-dimensional temperature data of the turbine blades, and obtain S2 aerodynamic data of the turbine blades through aerodynamic analysis. Using the highest physical rotational speed, three-dimensional temperature data of turbine blades, S2 aerodynamic data, and mechanical property data of turbine blade materials in each region as inputs, stress calculation and analysis are performed using a finite element model of turbine blades to obtain the first principal stress, second principal stress, third principal stress, and equivalent stress at the test location of the finite element model in the corresponding load step; based on the first principal stress, second principal stress, third principal stress, and equivalent stress at the test location in the corresponding load step, the stress correction stress at the test location in the corresponding load step is calculated. Using a standard round bar of turbine blade material as the test specimen, the creep rupture time of the test specimen was obtained under different test temperatures and test stresses. Using test temperature and test stress as inputs and creep rupture time as output, a unified equation for the comprehensive parameters of creep rupture and thermal strength of turbine blade material was established. Based on the corrected stress at the test location under the corresponding load step in each region, the equivalent stress at the test location in the corresponding region was obtained by analyzing the unified equation for the comprehensive parameters of creep rupture and thermal strength. Based on the equivalent stress and temperature at the turbine blade test location in each region, the creep fracture time at the test location is obtained by analyzing a unified equation that combines creep and thermal strength parameters. Based on the sustained fracture time at the test location, the linear cumulative damage at the test location of the turbine blade is calculated using the Miner linear cumulative damage model.

[0018] In this embodiment, the load spectrum during the flight climb phase of the actual flight mission is first divided into regions, and a finite element model of the turbine blade is established. The load spectrum from the actual flight climb process is applied to each region to calculate stress, thereby calculating the first principal stress, second principal stress, third principal stress, and equivalent stress at each node of the turbine blade finite element model. The stress at the test location is then corrected to obtain the equivalent stress at that location. The creep rupture time of the turbine blade is calculated using a unified equation combining creep and thermal strength parameters. Finally, Miner linear accumulation is used for damage calculation.

[0019] This damage assessment method considers the decrease in engine physical speed, aerodynamics, internal flow, and temperature due to fuel consumption during the climb phase of actual flight missions, thus improving the accuracy of load analysis. Furthermore, it corrects the stress at the turbine blade assessment location based on the first principal stress, second principal stress, third principal stress, and equivalent stress, facilitating the reasonable determination of the blade's critical location and improving stress calculation results. The equivalent stress in the equation unifying creep rupture and thermal strength parameters for calculating the sustained fracture time considers the influence of stress relaxation effects and is optimized using a multi-objective genetic optimization algorithm, unifying the fitting parameters at different temperatures and making the predicted linear cumulative damage results more accurate.

[0020] Example 2 See Figures 1 to 2A method for assessing persistent damage to turbine blades based on the load spectrum of a flight climb mission is presented, and the assessment process is as follows: Step 1: Divide the load spectrum of the flight climb phase to be analyzed into multiple regions, and extract the rotational speed and cumulative time of each region; In this embodiment, as shown... Figure 2 As shown, after determining the load spectrum of the climb phase to be analyzed in the actual flight profile of the engine, the climb phase is divided into stages with a step size of 1% of the initial physical rotational speed of the climb phase to be analyzed. One region is designated as the climb phase to be analyzed. A typical state. (The following is a possible interpretation / statement.) The highest physical speed within a region is taken as the stable physical speed within that region. Cumulative time within each region.

[0021] Step 2: Divide each region into multiple load steps. Using the overall performance data of the engine, air system data, turbine aerodynamic data, and material heat transfer performance data under each load step as boundary conditions, perform finite element heat transfer analysis of the turbine blades to obtain three-dimensional temperature data of the turbine blades. Obtain the S2 aerodynamic data of the turbine blades through aerodynamic analysis.

[0022] Step 3: Using the highest physical rotational speed of each region, turbine blade three-dimensional temperature data, S2 aerodynamic data, and turbine blade material mechanical property data as inputs, perform stress calculation and analysis using the turbine blade finite element model to obtain the first principal stress, second principal stress, third principal stress, and equivalent stress at the test location of the finite element model in the corresponding load step; based on the first principal stress, second principal stress, third principal stress, and equivalent stress at the test location in the corresponding load step, calculate the stress correction stress at the test location in the corresponding load step; the specific process is as follows: 2.1 Constructing a creep damage constitutive model for turbine blade materials, the creep damage constitutive model including a strain hardening model simulating the first stage of creep damage in turbine blade materials. And the Norton model for simulating the second stage of creep damage in turbine blade materials. ;in To determine the creep strain rate for the first stage of creep damage under the selected test temperature and stress, The test stress selected for the standard round bar of turbine blade material. The test temperature selected for the standard round bar of turbine blade material. To determine the creep strain in the first stage of creep damage under selected test temperature and stress, It is a natural constant. To determine the creep strain rate of the second stage of creep damage under selected test temperature and test stress, to These are the fitting parameters, obtained through data fitting; 2.2 Based on the tensile test results of the standard round bar specimens of turbine blade material, construct the stress-strain curve of the turbine blade material, input it into the finite element model of the turbine blade, and then input the fitting parameters of the creep damage constitutive model fitted in step 2.1. to ; 2.3 Using the highest physical rotational speed, turbine blade three-dimensional temperature data, S2 aerodynamic data, and turbine blade material mechanical property data for each region as input, finite element analysis is performed on the turbine blade to obtain the stress variation of the blade under each load step in each region with the cumulative working time. In this embodiment, the first region is taken as an example, and the stress is set according to the cumulative time of the first region. One load step. The load steps for other regions are the same as those for the first region.

[0023] 2.4 Based on the stress variation of the turbine blade in each region with the cumulative working time, the first principal stress, second principal stress, third principal stress and equivalent stress of all nodes of the turbine blade under each load step can be output through finite element analysis; 2.5 Corrected stress at the test location under the corresponding load step ;in, Indicates the assessment position is at the _____. i The first region j Under each load step t Correction stress at time, , m This represents the total number of regions during the climbing phase. , For the first i Total number of load steps in each region The assessment position is in the 1st place. i The first region j Under each load step t The first principal stress at time t, The assessment position is in the 1st place. i The first region j Under each load step t The second principal stress at time t, The assessment position is in the 1st place. i The first region j Under each load step t The third principal stress at time t, The assessment position is in the 1st place. i The first region j Under each load step t The equivalent force at time, The assessment position is in the 1st place. i The first region j Under each load stept Stress triaxiality at time t, These are the experience coefficients, ranging from 0 to 1; the assessment position is at the [number]th [position]. i The first region j Stress triaxiality under each load step .

[0024] Step 4: Using a standard round bar of turbine blade material as the test specimen, the creep rupture time of the test specimen is obtained under different test temperatures and test stresses. Using test temperature and test stress as inputs and creep rupture time as output, a unified equation for the comprehensive parameters of creep rupture and thermal strength of turbine blade material is established. Based on the corrected stress at the test location under the corresponding load step in each region, the equivalent stress at the test location in the corresponding region is obtained by analyzing the unified equation for the comprehensive parameters of creep rupture and thermal strength. In this embodiment, the equation establishing the unified comprehensive parameters of turbine blade material creep duration and thermal intensity is as follows: ,in, The sustained fracture stress of a standard round bar for turbine blade material at a selected test temperature. The test temperature selected for the standard round bar of turbine blade material. The time to sustained fracture of a standard round bar of turbine blade material at a selected test temperature and sustained fracture stress; , , , , , , , These are the model parameters, which are obtained by using a multi-objective genetic optimization algorithm with the goal of minimizing the preset loss function.

[0025] In this embodiment, a multi-objective genetic optimization algorithm is used, and a loss function is defined. The algorithm finds the minimum value of the loss function, adjusts the boundaries of the fitting parameters, the number of iterations, and the termination tolerance, and finally outputs the optimal fitting parameters. .in, The persistent breakage time obtained from multi-objective genetic optimization calculations.

[0026] In this embodiment, the equivalent stress at the test location is located in the corresponding region. ,in The assessment position is in the 1st place. i Equivalent stress in each region For the first i Cumulative time in the region For the first i The first in the region End time of load step Correction stress at time, , For the first i Number of load steps in the region For the first i The first in the region j The cumulative time of the load step. The assessment position is in the 1st place. i The power exponent of the region, , To analyze the turbine blade test position obtained by using a unified equation that combines creep and thermal stress parameters. i The first in the region x Corresponding temperature and corrected stress of the load step The duration of sustained fracture, Represents the differential symbol.

[0027] Step 5: Based on the equivalent stress and temperature at the turbine blade test location in each region, the creep rupture time at the test location is obtained by analyzing the equation that unifies the comprehensive parameters of creep rupture and thermal intensity. In this embodiment, after obtaining a unified equation expression for the comprehensive parameters of creep and thermal strength of the turbine blade material through data fitting, the equivalent stress and corresponding temperature at the test location of the turbine blade can be substituted into the equation expression to calculate the value at the test location. i The duration of sustained fracture in each region.

[0028] Step 6: Based on the sustained fracture time at the test location, the linear cumulative damage at the test location of the turbine blade is calculated using the Miner linear cumulative damage model. In this embodiment, the linear cumulative damage at the turbine blade testing location is described. ,in, For the first i Cumulative time in the region For the first i The turbine blade test location within the region is at the corresponding temperature equivalent stress. The cumulative time below, , m This represents the total number of regions during the climbing phase.

[0029] The above are merely preferred embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent substitutions, and improvements made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A method for assessing persistent damage to turbine blades based on the load spectrum of a flight climb mission, characterized in that, include: The load spectrum of the flight climb phase to be analyzed is divided into multiple regions, and the rotational speed and cumulative time of each region are extracted. Each region is divided into multiple load steps. The overall performance data of the engine, air system data, turbine aerodynamic data and material heat transfer performance data under each load step are used as boundary conditions to perform finite element heat transfer analysis of the turbine blades, obtain three-dimensional temperature data of the turbine blades, and obtain S2 aerodynamic data of the turbine blades through aerodynamic analysis. Using the highest physical rotational speed, three-dimensional temperature data of turbine blades, S2 aerodynamic data, and mechanical property data of turbine blade materials in each region as inputs, stress calculation and analysis are performed using a finite element model of turbine blades to obtain the first principal stress, second principal stress, third principal stress, and equivalent stress at the test location of the finite element model in the corresponding load step; based on the first principal stress, second principal stress, third principal stress, and equivalent stress at the test location in the corresponding load step, the stress correction stress at the test location in the corresponding load step is calculated. Using a standard round bar of turbine blade material as the test specimen, the creep rupture time of the test specimen was obtained under different test temperatures and test stresses. Using test temperature and test stress as inputs and creep rupture time as output, a unified equation for the comprehensive parameters of creep rupture and thermal strength of turbine blade material was established. The corrected stress at the test location in each region under the corresponding load step was also calculated. The equivalent stress at the test location in the corresponding region is obtained by analyzing the equation that unifies the creep duration and thermal strength parameters. Based on the equivalent stress and temperature at the turbine blade test location in each region, the creep fracture time at the test location is obtained by analyzing the equation that unifies the creep duration and thermal strength parameters. Based on the sustained fracture time at the test location, the linear cumulative damage at the test location of the turbine blade is calculated using the Miner linear cumulative damage model.

2. The method for assessing persistent damage to turbine blades according to claim 1, characterized in that, Using 1% of the initial physical rotational speed of the flight climb phase to be analyzed as a step, the load spectrum of the flight climb phase to be analyzed is divided into multiple regions.

3. The method for assessing persistent damage to turbine blades according to claim 1, characterized in that, Methods for obtaining the first principal stress, second principal stress, third principal stress, and equivalent stress at the corresponding load step in the finite element model assessment location include: A creep damage constitutive model for turbine blade materials is constructed, comprising a strain hardening model simulating the first stage of creep damage in turbine blade materials. And the Norton model for simulating the second stage of creep damage in turbine blade materials. ;in To determine the creep strain rate for the first stage of creep damage under the selected test temperature and stress, The test stress selected for the standard round bar of turbine blade material. The test temperature selected for the standard round bar of turbine blade material. To determine the creep strain in the first stage of creep damage under selected test temperature and stress, It is a natural constant. To determine the creep strain rate of the second stage of creep damage under selected test temperature and test stress, to These are the fitting parameters, obtained through data fitting; Based on the tensile test of a standard round bar for turbine blade materials, a true stress-strain curve for the turbine blade material is constructed. This true stress-strain curve, along with the parameters of the fitted creep damage constitutive model, is then analyzed. to Input the finite element model of the turbine blade; Using the highest physical rotational speed of each region, three-dimensional temperature data of turbine blades, S2 aerodynamic data, and mechanical property data of turbine blade materials as inputs, stress analysis is performed using a finite element model of turbine blades to obtain the first principal stress, second principal stress, third principal stress, and equivalent stress at the test location under the corresponding load step.

4. The method for assessing persistent damage to turbine blades according to claim 1, characterized in that, The corrected stress at the test location under the corresponding load step ;in, Indicates the assessment position is at the _____. i The first region j Under each load step t Correction stress at time, , m This represents the total number of regions during the climbing phase. , For the first i Total number of load steps in each region The assessment position is in the 1st place. i The first region j Under each load step t The first principal stress at time t, The assessment position is in the 1st place. i The first region j Under each load step t The second principal stress at time t, The assessment position is in the 1st place. i The first region j Under each load step t The third principal stress at time t, The assessment position is in the 1st place. i The first region j Under each load step t The equivalent force at time, The assessment position is in the 1st place. i The first region j Under each load step t Stress triaxiality at time t, These are empirical coefficients, with values ​​ranging from 0 to 1.

5. The method for assessing persistent damage to turbine blades according to claim 4, characterized in that, The assessment position is at i The first region j Stress triaxiality under each load step .

6. The method for assessing persistent damage to turbine blades according to claim 4, characterized in that, The equivalent stress at the test location in the corresponding region ,in The assessment position is in the 1st place. i Equivalent stress in each region For the first i Cumulative time in the region For the first i The first in the region End time of load step Correction stress at time, , For the first i Number of load steps in the region For the first i The first in the region j The cumulative time of the load step. The assessment position is in the 1st place. i The power exponent of the region, , To analyze the turbine blade test position obtained by using a unified equation that combines creep and thermal stress parameters. i The first in the region x Corresponding temperature and corrected stress of the load step The duration of sustained fracture, Represents the differential symbol.

7. The method for assessing persistent damage to turbine blades according to claim 1, characterized in that, The unified equation for the combined parameters of creep durability and thermal strength of turbine blade materials is as follows: ,in, The sustained fracture stress of a standard round bar for turbine blade material at a selected test temperature. The test temperature selected for the standard round bar of turbine blade material. The time to sustained fracture of a standard round bar of turbine blade material at a selected test temperature and sustained fracture stress; , , , , , , , These are the model parameters, which are obtained by using a multi-objective genetic optimization algorithm with the goal of minimizing the preset loss function.

8. The method for assessing persistent damage to turbine blades according to claim 1, characterized in that, Linear cumulative damage at the test location of turbine blades ,in, For the first i Cumulative time in the region For the first i The turbine blade test location within the region is at the corresponding temperature equivalent stress. The cumulative time below, , m This represents the total number of regions during the climbing phase.

Images