A design method for aeroengine turbine blade leading edge creep simulation component

Abstract

The present invention relates to a method for designing a leading edge creep simulation part of an aero-engine turbine blade, comprising the following steps: (1) obtaining a geometric model, operating conditions and material parameters of a turbine blade; (2) establishing a finite element creep analysis model of the turbine blade, calculating a leading edge creep dangerous part, determining a dangerous section, and obtaining creep strain and creep strain rate distribution data of the dangerous part; (3) extracting geometric features of the dangerous part as a basis for preliminary design of the shape of the simulation part, and preliminarily designing the simulation part with a uniform thickness curved plate specimen as the basic shape; and (4) correcting key dimensions of the simulation part in combination with the finite element creep analysis results of the preliminary design simulation part to ensure that the creep strain rate distribution of the simulation part is consistent with that of the dangerous part of the leading edge of the turbine blade, thereby completing the design of the simulation part.

Description

Technical Field

[0001] The invention belongs to the technical field of aerospace engines, and in particular relates to a design method for a leading edge creep simulation part of an aerospace engine turbine blade. Background Art

[0002] Turbine blades are important parts of aircraft engines, with dual complexity of structure and load. Turbine blades work under high temperature and high load for a long time, and bear great thermal load, centrifugal force and aerodynamic force. Due to the harsh working conditions, turbine blades become parts with frequent failures, and their life directly determines the service life of aircraft engines. Under the action of high temperature, centrifugal force and pressure load for a long time, although the stress level is lower than the yield limit, the material will still undergo continuous, slow and irreversible deformation, which will not only reduce the flight efficiency of the engine, but also cause radial elongation, torsion and bending of the blades. When the deformation increases to exceed the ductility range of the material, the blades will creep fracture. Among them, the leading edge of the turbine blade directly bears the impact of high-temperature and high-speed gas from the combustion chamber, which makes it bear large thermal load and aerodynamic load, and is prone to creep failure. Since the real blade test is expensive and difficult to carry out on a large scale, and the traditional creep life prediction based on the test results of standard parts is difficult to accurately predict the creep life prediction of turbine blades, it is of great significance to propose a design method for creep simulation parts of the leading edge of turbine blades.

[0003] The existing literature "Huo Junzhou, Sun Debin, Wu Hanyang, Wang Weizeng, Xue Lin. Multi-axis low-cycle creep / fatigue life prediction of high-pressure turbine blades based on new critical plane damage parameters [J]. Engineering Failure Analysis, 2019(106):104-159." proposed a creep-fatigue simulation component design method that uses the finite element linear elastic analysis method to determine the dangerous parts, and did not use the creep nonlinear constitutive relationship for finite element analysis. Summary of the invention

[0004] In order to overcome the shortcomings of the prior art, the present invention provides a design method for an aircraft engine turbine blade leading edge creep simulation part, which can fully reflect the creep strain rate and creep strain distribution of the turbine blade leading edge, and serve and support the creep life assessment of aircraft engine turbine blades.

[0005] The invention combines the results of turbine blade finite element creep simulation to design a simulation part that can reflect the characteristic structure, and realizes the simulation of the creep behavior of the real blade leading edge structure under laboratory conditions. The implementation steps of the invention are as follows:

[0006] The first step is to obtain the geometric model, working conditions and material parameters of the turbine blades; the geometric model refers to the three-dimensional geometric model of the turbine blades designed according to actual engineering requirements with the help of computer-aided design software; the working conditions include the speed and temperature field of the turbine blades under actual working conditions; the material parameters include the density of the turbine blade material, the elastic modulus, Poisson's ratio and creep constitutive model parameters at the working temperature. The typical working speed of the turbine blades is obtained through the statistics of the external engine working data; the working temperature field of the turbine blades is obtained through heat transfer analysis or actual measurement; the material performance parameters are obtained through material performance tests or material data manuals. For material parameters that change with temperature, such as elastic modulus and Poisson's ratio, the corresponding parameter values ​​within a certain range need to be obtained, and the temperature range taken needs to include the maximum and minimum values ​​in the turbine blade working temperature field; the creep model parameters are obtained by fitting the creep curve in the test or material manual.

[0007] In the second step, according to the geometric model, working conditions and material parameters of the first step, a finite element creep analysis model of the turbine blade is established, the leading edge creep danger zone is calculated, the dangerous section is determined, and the creep strain and creep strain rate distribution data of the dangerous section are obtained; the creep strain and creep strain rate distribution data include the maximum value of the creep strain rate of the dangerous section, the circumferential distribution of the creep strain rate along the inner wall of the dangerous section, and the creep strain distribution at different times; the creep strain and creep strain rate distribution are obtained by measuring the creep strain and creep strain rate of several points on the inner wall of the dangerous section.

[0008] The third step is to extract the geometric features of the dangerous parts as the basis for the preliminary design of the simulation based on the dangerous section position obtained in the second step and the geometric model of the turbine blade described in the first step, and to preliminarily design the simulation parts with the equal thickness curved plate as the basic shape; the geometric features include the leading edge thickness and the radius of curvature of the blade.

[0009] The fourth step is to design a turbine blade leading edge creep simulation part based on the creep strain and creep strain rate distribution data of the dangerous parts obtained in the second step, with four-point bending loading perpendicular to the length direction of the simulation part as the load form, establish a finite element analysis model of the turbine blade leading edge creep simulation part, perform creep simulation analysis on the model, adjust the key dimensions of the simulation part, so that the maximum value of the creep strain rate of the simulation part is equal to the maximum value of the creep strain rate of the dangerous section obtained in the second step; further, make the circumferential distribution of the creep strain rate along the inner wall of the simulation part and the creep strain distribution at different times consistent with the creep strain and creep strain rate distribution of the dangerous parts obtained in the second step, and complete the determination of the dimensional parameters of the simulation part and the load size of the uniaxial stretching of the simulation part; at this point, the design work of the turbine blade leading edge creep simulation part is completed.

[0010] The advantages of the present invention over the prior art are as follows: Based on the existing simulation part design method, the present invention proposes a method for designing a leading edge creep simulation part of an aero-engine turbine blade, which is used to study the influence of the leading edge of the blade on the creep life of the aero-engine turbine blade, and can simulate the creep test effect of the leading edge of the real turbine blade. The dangerous section is obtained by the present invention through nonlinear creep finite element simulation of the blade model; while reflecting the geometric characteristics of the leading edge of the turbine blade, the creep strain rate distribution is further made consistent with the dangerous section of the leading edge of the turbine blade, meeting the needs of the turbine blade leading edge creep test simulation, and serving and supporting the design of aero-engine turbine blade simulation parts and life prediction technology. At present, no relevant technology has been reported, and the present invention makes up for the shortcomings of related research. BRIEF DESCRIPTION OF THE DRAWINGS

[0011] Figure 1 A flowchart of the implementation of a method for designing a leading edge creep simulation component of an aircraft engine turbine blade according to the present invention;

[0012] Figure 2 It is a schematic diagram of a turbine blade geometric model and temperature field simulation results of the present invention;

[0013] Figure 3 It is a schematic diagram of the dangerous section and basic dimensions of the present invention;

[0014] Figure 4 The initial shape design result of the simulation part of the present invention;

[0015] Figure 5 The optimized shape and geometric dimensions of the simulation part of the present invention;

[0016] Figure 6 The creep strain rate distribution along the axis direction of the turbine blade of the present invention and the simulation part is compared;

[0017] Figure 7 The creep strain distribution of the turbine blade of the present invention and the simulation part along the axial direction at 100, 200 and 300 hours is compared. DETAILED DESCRIPTION

[0018] In order to make the purpose, technical solutions and advantages of the present invention more clearly understood, the present invention is further described in detail below in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not intended to limit the present invention. In addition, the technical features involved in the various embodiments of the present invention described below can be combined with each other as long as they do not conflict with each other.

[0019] The technical solution of a method for designing a leading edge creep simulation component of an aircraft engine turbine blade of the present invention is further described below in conjunction with the accompanying drawings.

[0020] like Figure 1 As shown, a method for designing a leading edge creep simulation component of an aircraft engine turbine blade of the present invention is implemented in the following steps:

[0021] The first step is to obtain the geometric model, operating conditions and material parameters of the actual turbine blades; the geometric model refers to the three-dimensional geometric model of the turbine blades designed according to actual engineering requirements with the help of computer-aided design software; the operating conditions include the rotational speed and temperature field of the turbine blades under actual working conditions; the material parameters include the density of the turbine blade material, the elastic modulus, Poisson's ratio and creep constitutive model parameters at the working temperature. For example, the takeoff state of a certain type of aircraft engine turbine blade is analyzed. The temperature range of the temperature field of the turbine blade in the takeoff state is 525℃-1068℃, the rotation speed is 45450rpm, and the blade geometry model and temperature field simulation results are as follows: Figure 2 As shown in the figure, the blade material is DD6 single crystal nickel-based high temperature alloy. According to the China Aviation Materials Handbook, the density is 8780kg / m 3 The elastic modulus that varies with temperature is shown in Table 1, and the Poisson's ratio that varies with temperature is shown in Table 2. The creep constitutive model adopts the Norton model, and the expression is as follows:

[0022]

[0023] in is the creep strain rate, C1, C2, C3 are constants related to materials and temperature, σ is the equivalent stress, and T is the service temperature. The creep curve of the material is fitted in the material manual to obtain the Norton creep model parameters C1, C2, C3 at different temperatures, and the creep model parameters are interpolated to obtain the creep model parameters within the service temperature range.

[0024] Table 1 Elastic modulus of single crystal nickel-based superalloy at different temperatures and orientations

[0025]

[0026] Table 2 Poisson's ratio of single crystal nickel-based superalloy at different temperatures and orientations

[0027]

[0028] In the second step, according to the geometric model, working conditions and material parameters of the first step, the actual blade finite element creep analysis model is established, the leading edge creep danger position is calculated, the dangerous section is determined, and the creep strain and creep strain rate distribution data of the dangerous position are obtained; the creep strain and creep strain rate distribution data include the maximum value of the creep strain rate of the dangerous section, the circumferential distribution of the creep strain rate along the inner wall of the dangerous section, and the creep strain distribution at different times. In this example, the dangerous part of the leading edge of the blade is the chamfer of the rib plate; the height of the dangerous section position is 84.963mm; the creep strain rate is obtained by time differentiation of the strain, and the creep strain rate distribution and the creep strain distribution of 100h, 200h, and 300h are extracted.

[0029] In the third step, based on the dangerous section position obtained in the second step and the turbine blade model described in the first step, the geometric features of the dangerous part are extracted as the basis for the preliminary design of the simulation, and the simulation part is preliminarily designed with the uniform thickness curved plate as the basic shape; the geometric features include the leading edge thickness and curvature radius of the blade. In this example, the section of the dangerous part of the leading edge is as follows Figure 3 As shown; the leading edge thickness of the blade is 0.7mm, the radius of curvature is 6mm, and the preliminary configuration of the design simulation part is as follows Figure 4 As shown, its thickness and radius of curvature are the same as those of an actual blade.

[0030] The fourth step is to design a turbine blade leading edge creep simulation part based on the creep strain and creep strain rate distribution data of the dangerous parts obtained in the second step, with four-point bending loading perpendicular to the length direction of the simulation part as the load form, establish a finite element analysis model of the turbine blade leading edge creep simulation part, perform creep simulation analysis on the model, adjust the key dimensions of the simulation part, and make the maximum value of the creep strain rate of the simulation part equal to the maximum value of the creep strain rate of the dangerous section obtained in the second step; further, make the circumferential distribution of the creep strain rate along the inner wall of the simulation part and the creep strain distribution at different times consistent with the creep strain and creep strain rate distribution of the dangerous parts obtained in the second step, and complete the determination of the dimensional parameters of the simulation part and the load size of the uniaxial tension of the simulation part. In this example, firstly, a finite element creep analysis is performed on the preliminary design results of the simulation part according to the actual loading conditions. Since the creep strain rate of the initial design results of the simulation part coincides with the maximum value of the creep strain rate of the real blade, but the circumferential distribution of the creep strain rate does not match, a semicircular arc is added in the test section to change the circumferential gradient distribution result of the creep strain rate of the inner wall of the simulation part. Taking the circumferential distribution of creep strain rate of the inner wall of the leading edge of a real blade as the object, by adjusting the four-point bending loading position a, b, the curvature of the curved plate θ', and the radius of the semicircular arc r', as shown in Figure 5As shown in the figure, the distribution of the damage control parameter (i.e. creep strain rate) of the simulated component within the critical distance range is the same as that of the real component. By optimizing the parameters, the final size distribution of the simulated component test section is obtained, where a is 15mm, b is 50mm, θ' is 60°, and r' is 0.4mm. At this time, the creep strain rate distribution of the leading edge simulated component and the real blade is compared. Figure 6 The creep strain distribution at different times is shown in Figure 7 As shown in the figure, within the critical distance range, the maximum error does not exceed 5.7%, thus completing the design of the simulation part.

[0031] The above embodiments are provided only for the purpose of describing the present invention, and are not intended to limit the scope of the present invention. The scope of the present invention is defined by the appended claims. Various equivalent substitutions and modifications made without departing from the spirit and principles of the present invention should all be included within the scope of the present invention.

Claims

1. A method for designing a creep simulation component of a leading edge of an aircraft engine turbine blade, characterized in that: The steps include: Step (1): obtaining a geometric model, operating conditions and material parameters of a turbine blade; the geometric model refers to a three-dimensional geometric model of a turbine blade designed according to actual engineering requirements with the aid of computer-aided design software; the operating conditions include the rotation speed and temperature field of the turbine blade under actual working conditions; the material parameters include the density of the turbine blade material, the elastic modulus at the working temperature, the Poisson's ratio and the creep constitutive model parameters; Step (2): according to the geometric model, working conditions and material parameters of step (1), a finite element creep analysis model of the turbine blade is established, the leading edge creep danger position is calculated, the dangerous section is determined, and the creep strain and creep strain rate distribution data of the dangerous part are obtained; the creep strain and creep strain rate distribution data include the maximum value of the creep strain rate of the dangerous section, the circumferential distribution of the creep strain rate along the inner wall of the dangerous section, and the creep strain distribution at different times; Step (3): Based on the dangerous section position obtained in step (2) and the geometric model of the turbine blade in step (1), the geometric features of the dangerous part are extracted as the basis for the preliminary design of the simulation part, and the simulation part is preliminarily designed with the uniform thickness curved plate as the basic shape; the geometric features include the leading edge thickness and the curvature radius of the blade; Step (4): Based on the creep strain and creep strain rate distribution data of the dangerous parts obtained in step (2), a turbine blade leading edge creep simulation part is designed with four-point bending loading perpendicular to the length direction of the simulation part as the load form, a finite element analysis model of the turbine blade leading edge creep simulation part is established, a creep simulation analysis is performed on the model, and the key dimensions of the simulation part are adjusted so that the maximum value of the creep strain rate of the simulation part is equal to the maximum value of the creep strain rate of the dangerous section obtained in step (2), and the circumferential distribution of the creep strain rate along the inner wall of the simulation part is consistent with the creep strain rate distribution of the dangerous parts obtained in step (2), thereby completing the determination of the dimensional parameters of the simulation part and the load size of the uniaxial tension of the simulation part.

2. The method for designing a leading edge creep simulation component of an aircraft engine turbine blade according to claim 1, characterized in that: In the step (2), the creep critical section of the turbine blade is determined by performing finite element creep simulation analysis on the actual blade model.

3. The method for designing a leading edge creep simulation component of an aircraft engine turbine blade according to claim 1, characterized in that: In the step (4), the simulation component is loaded by four-point bending loading perpendicular to the length direction of the simulation component.

4. The method for designing a leading edge creep simulation component of an aircraft engine turbine blade according to claim 1, characterized in that: In the step (4), the basic shape of the simulation part is a curved plate of equal thickness.

5. The method for designing a leading edge creep simulation component of an aircraft engine turbine blade according to claim 1, characterized in that: In the step (4), the method for determining the critical dimension value is that the maximum error between the creep strain rate distribution of the simulated component and the creep strain rate distribution of the real structure within the critical distance is no more than 5.7%.

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