A method for designing a creep simulator of an aeroengine turbine blade trailing edge cooling hole
By designing a creep simulation component for the trailing edge cooling holes of aero-engine turbine blades, the problem of the lack of simulation component design in the existing technology has been solved, realizing low-cost creep test simulation and life prediction, and meeting the creep analysis requirements of turbine blades.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- BEIHANG UNIV
- Filing Date
- 2022-07-26
- Publication Date
- 2026-07-31
AI Technical Summary
Existing technologies lack effective simulation design methods to study the impact of turbine blade trailing edge cooling holes on creep life, making it difficult to conduct high-cost real blade tests and lacking reliable creep prediction support.
A creep simulation component for the trailing edge cooling hole of an aero-engine turbine blade was designed. By determining the geometric dimensions and material property parameters of the simulation component, a finite element creep analysis model was established, and the position and shape of the hole in the simulation component were optimized to simulate the creep strain rate and stress distribution of a real turbine blade.
It enables creep testing of the trailing edge cooling hole area of a real turbine blade at a relatively low cost, reflects the geometric characteristics of the trailing edge cooling hole of the turbine blade, meets the requirements of creep testing, and supports the design of simulated turbine blades and life prediction.
Smart Images
Figure CN115329482B_ABST
Abstract
Description
Technical Field
[0001] This invention belongs to the field of aerospace engine technology, specifically relating to a design method for a creep simulation component of a cooling hole at the trailing edge of an aero-engine turbine blade. Background Technology
[0002] Turbine blades are critical safety components of aero-engines, enduring complex alternating force and temperature loads during service, thus limiting the engine's lifespan. Regarding durability, turbine blades are subjected to immense centrifugal loads from the high-speed turbine rotation, aerodynamic pressure loads from high-speed exhaust gas impacting the blades, and temperature loads from high-temperature exhaust gas and temperature gradients. Under prolonged exposure to high temperatures and centrifugal and pressure loads, even with stress levels below the yield limit, the material undergoes continuous, slow, and irreversible deformation. This not only reduces engine flight efficiency but also causes radial elongation, torsion, and bending of the blades. Furthermore, as the deformation exceeds the material's ductility, creep fracture occurs. Currently, turbine blade creep life prediction lacks experimental data support, necessitating a complete, reliable, and experimentally validated engine strength design system. Cooling holes are a common structure on turbine blades, but their relatively small size can cause stress concentration, significantly reducing creep life. Therefore, it is necessary to analyze the creep performance of cooling hole areas. However, actual blade testing is costly, and there is a lack of simulation design methods for cooling hole areas; therefore, a corresponding simulation design method is needed. Summary of the Invention
[0003] To overcome the shortcomings of existing technologies, this invention provides a method for designing a creep simulation component for cooling holes at the trailing edge of aero-engine turbine blades, specifically including the following steps:
[0004] The solution of this invention is: a creep simulation design method for trailing edge cooling holes of aero-engine turbine blades. Taking the trailing edge cooling holes of the turbine blade as the simulation area, the simulation component is determined to be a flat plate of uniform thickness based on the geometric dimensions of the trailing edge of the turbine blade's critical section and the trailing edge cooling holes. The length of the plate is the length of the trailing edge, and the thickness of the plate is the maximum thickness of the trailing edge. The holes in the simulation component are cylindrical through holes along the length direction. By optimizing the position of the holes in the simulation component, the position of the point of maximum creep strain rate, the creep strain rate, and the Von Mises equivalent stress are ensured to be consistent with the trailing edge cooling holes of the turbine blade's critical section. The implementation steps are as follows:
[0005] Step (1): Determine the critical section of the turbine blade, obtain the geometric model of the critical section, and measure the geometric dimensions of the trailing edge and the cooling hole at the trailing edge of the critical section;
[0006] Step (2): Obtain the operating conditions of the turbine blade and the material performance parameters under the operating conditions; the operating conditions include the operating temperature of the turbine blade, the centrifugal load and the aerodynamic load it is subjected to; the material performance parameters include the density of the turbine blade material, the elastic modulus, Poisson's ratio, expansion coefficient and creep constitutive parameters at the operating temperature;
[0007] Step (3): Based on the geometric model in step (1) and the working conditions in step (2) and the material performance parameters under the working conditions, establish a finite element creep analysis model for the turbine blade, calculate and obtain the point with the maximum creep strain rate at the trailing edge of the critical section cooling hole, and obtain the creep strain rate and creep strain distribution at different times at this point.
[0008] Step (4): Determine the basic shape of the simulated part with holes based on the geometry of the trailing edge and the trailing edge cooling hole; According to the geometry described in step (1), determine that the thickness change from the trailing edge of the critical section to the blade body is much smaller than the maximum thickness of the trailing edge, and the thickness of the trailing edge is much smaller than the length of the trailing edge; Determine that the trailing edge cooling hole is a cylindrical through hole along the length of the trailing edge, and forms a certain angle with the cross section of the turbine blade, and that the angle and diameter of the hole in the simulated part are the same as the angle and diameter of the trailing edge cooling hole, thus completing the determination of the basic shape of the simulated part;
[0009] Step (5): By adjusting the position of the simulated hole in the thickness direction and the uniaxial tensile load described in step (4), finite element creep analysis is performed to ensure that the position of the maximum creep strain rate point, the creep strain rate distribution, and the creep strain distribution at different times of the simulated hole are the same as the position of the maximum creep strain rate point, the creep strain rate distribution, and the creep strain distribution at different times of the cooling hole at the trailing edge of the dangerous section of the turbine blade.
[0010] Furthermore, the critical section of the turbine blade in step (1) is determined by finite element simulation or experiment.
[0011] Furthermore, the geometric model refers to the three-dimensional geometric model of the turbine blade designed using computer-aided design software.
[0012] Furthermore, the geometric dimensions of the trailing edge and the trailing edge cooling hole include the thickness of the trailing edge of the critical section, the thickness that varies between the trailing edge and the blade body, the length of the trailing edge, the diameter of the trailing edge cooling hole, the length of the trailing edge cooling hole, and the spatial position of the trailing edge cooling hole relative to the trailing edge.
[0013] Furthermore, in step (4), the simulation component is a flat plate of uniform thickness, the length of which is the length of the trailing edge and the thickness is the maximum thickness of the trailing edge.
[0014] The advantages of this invention compared to existing technologies are as follows: Based on existing simulation design methods, this invention proposes a creep simulation design method for trailing-edge cooling holes on aero-engine turbine blades. This method is used to study the impact of trailing-edge cooling holes on the creep life of aero-engine turbine blades and can simulate the creep test effects of the trailing-edge cooling hole area on real turbine blades. While reflecting the geometric characteristics of the trailing-edge cooling hole area, this invention further ensures that the location of the maximum creep strain rate, the creep strain rate distribution, and the creep strain distribution at different times are consistent with the trailing-edge cooling hole at the critical section of the turbine blade. This meets the needs of creep test simulation for trailing-edge cooling holes on turbine blades and serves and supports the design and life prediction technology of aero-engine turbine blade simulation components. Currently, no related technologies have been reported, and this invention fills a gap in related research. Attached Figure Description
[0015] Figure 1 The present invention provides a design process for a creep simulation component of a cooling hole at the trailing edge of an aero-engine turbine blade.
[0016] Figure 2 This is a schematic diagram of the trailing edge of a critical section of a turbine blade.
[0017] Figure 3 The curves show the change of creep strain over time under different temperatures and loads;
[0018] Figure 4 The initial shape and geometry of the simulated part;
[0019] Figure 5 The optimized shape and geometry of the simulated part;
[0020] Figure 6 A comparison of the axial distribution of creep strain rate at the bore edge of turbine blades and simulated parts;
[0021] Figure 7 A comparison of the axial distribution of creep strain along the bore edge of turbine blades and simulated parts at 100, 200, and 300 hours. Detailed Implementation
[0022] To make the objectives, technical solutions, and advantages of this invention clearer, the invention will be further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative and not intended to limit the invention. Furthermore, the technical features involved in the various embodiments of this invention described below can be combined with each other as long as they do not conflict with each other.
[0023] The following description, in conjunction with the accompanying drawings, further illustrates the technical solution of the creep simulation component design method for the trailing edge cooling hole of an aero-engine turbine blade according to the present invention.
[0024] like Figure 1 As shown, this invention relates to a method for designing a creep simulation component for a cooling hole at the trailing edge of an aero-engine turbine blade, and the implementation steps are as follows:
[0025] The first step is to determine the critical section of the turbine blade, obtain its geometric model, and measure the geometric dimensions of the trailing edge and trailing edge cooling holes of the critical section. The critical section of the turbine blade is determined through finite element simulation or experimentation. The geometric model refers to the three-dimensional geometric model of the turbine blade designed using computer-aided design software. The dimensions of the trailing edge and trailing edge cooling holes include the thickness of the trailing edge of the critical section, the thickness variation between the trailing edge and the blade body, the length of the trailing edge, the diameter of the trailing edge cooling holes, the length of the trailing edge cooling holes, and the spatial position of the trailing edge cooling holes relative to the trailing edge. For example, as... Figure 2 The trailing edge portion of the critical section of a turbine blade shown has a thickness that gradually increases from 0.8 mm to 1.8 mm from the edge to the blade body. The length of the trailing edge is 5.9 mm, the diameter of the cooling hole at the trailing edge is 0.6 mm, the length of the cooling hole is 6 mm, and the angle between the cooling hole and the trailing edge section is 10°.
[0026] The second step involves obtaining the operating conditions of the turbine blades and the material performance parameters under those conditions. The operating conditions include the operating temperature of the turbine blades, and the centrifugal and aerodynamic loads they are subjected to. The material performance parameters include the density of the turbine blade material, and its elastic modulus, Poisson's ratio, and creep constitutive parameters at the operating temperature. An analysis of the turbine blades' takeoff state is conducted. The temperature range of the turbine blade's takeoff temperature field is 525℃-1068℃, the rotational speed is 45450rpm, and the material used is DD6 single-crystal nickel-based superalloy with a density of 8.78g / cm³ according to the Chinese Aviation Materials Handbook. 3 The elastic modulus as a function of temperature is shown in Table 1, and the Poisson's ratio as a function of temperature is shown in Table 2. The Norton creep constitutive model is adopted, based on... Figure 3 The parameters of the Norton creep constitutive material are obtained by fitting the creep strain versus time curves at different temperatures.
[0027] Table 1 Elastic modulus of single-crystal nickel-based superalloys at different temperatures and orientations
[0028]
[0029] Table 2 Poisson's ratios of single-crystal nickel-based superalloys at different temperatures and orientations.
[0030]
[0031] The third step involves establishing a finite element creep analysis model for the turbine blade based on the geometric model from the first step, the operating conditions from the second step, and the material performance parameters under the operating conditions. This model calculates the point with the maximum creep strain rate at the trailing edge of the critical section cooling hole and obtains the creep strain rate distribution along the axial direction and the creep strain distribution at different times. The creep strain rate is obtained by differentiating the strain over time. The node coordinates of the point with the maximum strain rate at the hole edge are calculated, and the strain rate distribution and the creep strain distribution at 100h, 200h, and 300h are extracted.
[0032] The fourth step is to determine the basic shape of the simulated component with holes based on the trailing edge and the dimensions of the trailing edge cooling holes. Based on the geometric dimensions from the first step, it is determined that the thickness variation from the trailing edge to the blade at the critical section is much smaller than the maximum thickness of the trailing edge, and the thickness of the trailing edge is much smaller than its length. The simulated component uses a flat plate of uniform thickness, with the length of the plate being the length of the trailing edge and the thickness of the plate being the maximum thickness of the trailing edge. The trailing edge cooling holes are determined to be cylindrical through-holes along the length of the trailing edge, forming a certain angle with the blade cross-section. The holes in the simulated component are also cylindrical through-holes along their length, and the angle and diameter of the holes are the same as those of the trailing edge cooling holes. This completes the determination of the basic shape of the simulated component. Figure 4 As shown, the test section plate of the simulation part has a width L of 5.9 mm, a thickness of 1.8 mm, a hole diameter d of 0.6 mm, an angle θ between the hole and the edge of the plate of 80°, and a hole length of 6 mm.
[0033] The fifth step involves adjusting the position of the simulated hole in the thickness direction and the shape and dimensions of the plate in the fourth step, followed by finite element creep analysis. This ensures that the location of the maximum creep strain rate at the hole edge, the creep strain rate distribution, and the creep strain distribution at different times are identical to those of the cooling hole at the trailing edge of the turbine blade's critical section. First, finite element creep analysis is performed on the simulated hole at the thickness center and near the plate edge to obtain the location of the maximum creep strain rate, the creep strain rate distribution, and the creep strain distribution at different times. Since the creep strain rate gradient does not meet the requirements, a circular arc notch is added near the plate edge of the hole. The radius r and depth c of the circular arc notch are optimized, as follows... Figure 5 As shown. When the radius r is 4.5 mm and the depth c is 0.5 mm, the location of the point of maximum creep strain rate, the creep strain rate distribution, and the creep strain distribution at different times of the simulated part are consistent with those of the blade. The creep strain rate distribution is as follows. Figure 6 As shown, the creep strain distribution at different times is as follows: Figure 7 As shown, the design of the simulation component is now complete.
[0034] The above embodiments are provided merely for the purpose of describing the present invention and are not intended to limit the scope of the invention. The scope of the invention is defined by the appended claims. Various equivalent substitutions and modifications made without departing from the spirit and principles of the invention should be covered within the scope of the invention.
Claims
1. A method for designing a creep simulation component for cooling holes on the trailing edge of an aero-engine turbine blade, characterized in that, Includes the following steps: Step (1): Determine the critical section of the turbine blade, obtain the geometric model of the critical section, and measure the geometric dimensions of the trailing edge and the cooling hole at the trailing edge of the critical section; Step (2): Obtain the operating conditions of the turbine blade and the material performance parameters under the operating conditions; the operating conditions include the operating temperature of the turbine blade, the centrifugal load and the aerodynamic load it is subjected to; the material performance parameters include the density of the turbine blade material, the elastic modulus, Poisson's ratio, expansion coefficient and creep constitutive parameters at the operating temperature; Step (3): Based on the geometric model in step (1) and the working conditions in step (2) and the material performance parameters under the working conditions, establish a finite element creep analysis model for the turbine blade, calculate and obtain the point with the maximum creep strain rate at the trailing edge of the critical section cooling hole, and obtain the creep strain rate and creep strain distribution at different times at this point. Step (4): Determine the basic shape of the simulated part with holes based on the geometry of the trailing edge and the trailing edge cooling holes; Based on the geometric dimensions described in step (1), it is determined that the thickness variation from the trailing edge of the critical section to the blade body is less than the maximum thickness of the trailing edge, and the thickness of the trailing edge is less than the length of the trailing edge; the thickness of the trailing edge gradually increases from 0.8 mm to 1.8 mm from the edge to the blade body, and the length of the trailing edge is 5.9 mm; the width L of the test section plate of the simulation part is 5.9 mm, and the thickness is 1.8 mm; it is determined that the cooling hole at the trailing edge is a cylindrical through hole along the length direction of the trailing edge, and forms an angle of 10° with the cross-section of the turbine blade, and the angle and diameter of the hole in the simulation part are the same as the angle and diameter of the cooling hole at the trailing edge, thus completing the determination of the basic shape of the simulation part; Step (5): By adjusting the position of the simulated hole in the thickness direction and the uniaxial tensile load described in step (4), finite element creep analysis is performed to ensure that the position of the maximum creep strain rate point at the hole edge of the simulated hole, the creep strain rate distribution, and the creep strain distribution at different times are the same as the position of the maximum creep strain rate point at the hole edge of the cooling hole at the trailing edge of the dangerous section of the turbine blade, the creep strain rate distribution, and the creep strain distribution at different times.
2. The method for designing a creep simulation component for a cooling hole on the trailing edge of an aero-engine turbine blade according to claim 1, characterized in that: The critical section of the turbine blade in step (1) is determined by finite element simulation or experiment.
3. The method for designing a creep simulation component for a cooling hole on the trailing edge of an aero-engine turbine blade according to claim 1, characterized in that: The geometric model refers to the three-dimensional geometric model of the turbine blade designed using computer-aided design software.
4. The method for designing a creep simulation component for a cooling hole on the trailing edge of an aero-engine turbine blade according to claim 1, characterized in that: The geometry of the trailing edge and the trailing edge cooling hole includes the thickness of the trailing edge at the critical section, the thickness that varies between the trailing edge and the blade body, the length of the trailing edge, the diameter of the trailing edge cooling hole, the length of the trailing edge cooling hole, and the spatial position of the trailing edge cooling hole relative to the trailing edge.
5. The method for designing a creep simulation component for a trailing edge cooling hole of an aero-engine turbine blade according to claim 1, characterized in that: In step (4), the simulation component is a flat plate of equal thickness, the length of which is the length of the trailing edge and the thickness is the maximum thickness of the trailing edge.