A method for designing a turbine blade mockup dedicated to OEI based on full-field stress gradients

Abstract

The present application belongs to the technical field of aero-engines, and particularly relates to a design method of a turbine blade simulation piece special for OEI based on a full-field stress gradient, which can verify whether a space engine blade fails at a dangerous rotating speed after experiencing OEI based on an existing simulation piece design method by considering the stress gradient distribution of the whole blade body, introduce a safety reserve coefficient, define a dangerous cross section, and extract the temperature and stress distribution under the dangerous cross section by taking the whole cross section as a special dangerous surface, so as to serve and support the design of an aero-engine turbine blade simulation piece and the static strength evaluation and fracture rotating speed prediction of the blade, and make up for the insufficient research in the field.

Description

Technical Field

[0001] This invention belongs to the field of aero-engine technology, specifically relating to a design method for OEI-specific turbine blade simulation components based on full-field stress gradient. Background Technology

[0002] One-engine inoperative (OEI) power state refers to the engine shutdown phenomenon that occurs in twin-engine or multi-engine aircraft during in-flight operations due to common causes such as poor fuel or air intake, mechanical failure, cloud icing, and improper use of anti-icing systems. Rotary aircraft typically take off and land vertically without sufficient runways or helicopter landing pads. When a single-engine failure occurs during critical flight phases requiring high power, such as takeoff or landing, the remaining engine must apply a short-term high-power emergency power state to ensure sufficient power for a safe landing. Turbine blades are one of the core components of aero engines, typically operating in complex high-temperature, high-pressure environments. They are subjected to a combination of centrifugal, aerodynamic, and thermal loads under high-speed rotation. During OEI, the sudden overheating and overload can easily cause material damage, and in severe cases, blade failure and fracture can occur during OEI. Furthermore, due to the complex geometry of real blades, which often have tilted and torsional structures and geometric discontinuities such as flanges, extension roots, and tenons, and transmit torque to the rotor disk, they are subjected to combined tensile, bending, and torsional loads for extended periods. On the one hand, the actual blade structure is complex, difficult to manufacture, and costly to process; on the other hand, structural-level composite loads are difficult to apply. Therefore, in practical engineering, simulated components can be used to replace the actual structure for verification tests. When designing the simulated component, firstly, standard tensile test specimens with uniform axial cross-sections cannot reflect the influence of the axial stress gradient of the turbine blade on the critical section. Secondly, when the simulated component is designed to scale down proportionally to the turbine blade shape, its curved surface machining increases the difficulty and cost of processing. When the stress gradient at the critical section is not large, the shape of the critical section has a negligible impact on tensile strength. In this case, the simulated component can be designed as a specimen with a circular cross-section and a variable axial cross-section. Simultaneously, since the OEI state involves short-term overheating and overload, the temperature gradient caused by short-term overheating of the actual blade structure can be incorporated into the simulated component design by designing a variable cross-section. Summary of the Invention

[0003] This invention provides a design method for aero-engine turbine blade simulation components based on full-field stress gradient. It leverages existing simulation component design methods by considering the full-field stress gradient distribution across the blade body to verify whether aero-engine blades fail at critical speeds. Simultaneously, it introduces a safety reserve factor, defines a critical section, and extracts the temperature and stress distribution under the critical section as a special critical surface, serving as the simulation component's evaluation section. This method reflects the geometric characteristics of single-crystal turbine blades, ensuring that the characteristic stress values ​​of the simulated critical section are consistent with those of the real blade's critical section. Therefore, it can serve and support the design of aero-engine turbine blade simulation components, blade static strength assessment, and fracture speed prediction, thus addressing the shortcomings in related research in this field.

[0004] A design method for a dedicated turbine blade simulator based on full-field stress gradient using OEI (Optical Evolution of the Inertial Measurement) includes the following steps:

[0005] S1: Obtain the geometric model, operating conditions, and material parameters of the turbine blade;

[0006] S2: Based on the geometric model, operating conditions, and material parameters of the turbine blade, establish a finite element analysis model of the turbine blade, calculate the safety reserve coefficient of each section, and determine the critical section; obtain the elastic modulus and Poisson's ratio data of the critical part, as well as the normal stress distribution of the critical section, and calculate the normal tensile force F borne by the critical section;

[0007] S3: Based on the determined location of the critical section, extract the temperature of the critical section and the safety reserve factor of the section, and perform equivalent stress on different sections of the blade at the temperature under the critical section to obtain a three-dimensional model of the simulated test section.

[0008] S31: Based on the same safety reserve factor and temperature, and combined with the stress gradient distribution of the blade, the material strength change caused by different blade temperature distribution is used as the basis for design consideration of the simulation part. The characteristic stress of the three-dimensional model of the simulation part test section is determined, and then the geometric dimensions of the three-dimensional model of the simulation part test section are determined.

[0009] S32: Perform finite element analysis on the simulated test section to determine the shape of the three-dimensional model of the simulated test section;

[0010] S33: Lofting out each section of the simulated part to obtain a three-dimensional model of the test section of the simulated part;

[0011] S4: Based on the three-dimensional model of the simulated test section obtained by lofting, clamping ends that conform to the tensile test specimen standard are designed at both ends of the simulated part, and it is ensured that the characteristic stress of the simulated part's critical section is consistent with that of the turbine blade's critical section under the action of normal tensile force F at the critical section.

[0012] Based on existing simulation design methods, this method can verify whether aerospace engine blades fail at critical speeds by considering the stress gradient distribution in the blade body. Simultaneously, a safety reserve factor is introduced, a critical section is defined, and the entire section is used as a special critical surface to extract the temperature and stress distribution under the critical section, serving as the simulation test section. This method reflects the geometric characteristics of single-crystal turbine blades, ensuring that the characteristic stress values ​​of the simulated critical section are consistent with those of the real blade's critical section. Furthermore, it can serve and support the design of aerospace engine turbine blade simulation components, blade static strength assessment, and fracture speed prediction, thus filling gaps in related research in this field.

[0013] Furthermore, in S1, the geometric model refers to a three-dimensional geometric model of a turbine blade designed and completed according to actual engineering requirements using computer-aided design software.

[0014] Furthermore, in S1, the operating conditions include the rotational speed and temperature field of the turbine blade under actual working conditions; wherein, the operating rotational speed of the turbine blade is obtained through statistical analysis of the operating data of the external engine, and the operating temperature field of the turbine blade is obtained through heat transfer analysis or actual measurement.

[0015] Furthermore, in S1, the material parameters include the density of the turbine blade material, the elastic modulus at the operating temperature, and the Poisson's ratio; wherein, the material parameters are obtained through material performance tests or material data sheets.

[0016] Furthermore, in S2, the safety reserve factor refers to the ratio of the yield strength of the material in the

[001] direction at that temperature to the simulated normal stress, and its calculation expression is:

[0017]

[0018] In the formula, n is the safety reserve coefficient, σ 0.2 S represents the yield strength of the material in the

[001] direction. 11 The normal stress is obtained through simulation to align with the material orientation.

[0019] Furthermore, in step S2, the method for determining the critical section specifically includes the following steps:

[0020] S21: Based on the finite element analysis model of the turbine blade, the stress and temperature distribution of each section of the turbine blade are obtained, and the safety reserve factor of each section is calculated.

[0021] S22: Based on the simulation results and the calculated safety reserve coefficient, the danger point is obtained, and the cross section where the danger point is located is determined as the dangerous section of the blade.

[0022] Furthermore, in S2, the calculation expression for the normal tensile force F borne by the critical section is:

[0023]

[0024] In the formula, σ is the average stress at the critical section; s is the area of ​​the critical section.

[0025] Furthermore, in S3, the calculation expression for the other cross-sectional areas is as follows:

[0026]

[0027] In the formula, s' is the area of ​​other cross sections, n′ is the safety reserve factor of other cross sections, n0 is the safety reserve factor of the dangerous cross section, and s is the area of ​​the dangerous cross section.

[0028] The beneficial effects of this invention are as follows:

[0029] This invention, based on existing simulation design methods, verifies whether aerospace engine blades fail at critical speeds by considering the stress gradient distribution across the entire blade surface. It also introduces a safety reserve factor, defines a critical section, and extracts the temperature and stress distribution under the critical section as a special critical surface for use as the simulation test section. By reflecting the geometric characteristics of single-crystal turbine blades and ensuring that the characteristic stress values ​​of the simulated critical section are consistent with those of the real blade's critical section, this invention can serve and support the design of aerospace engine turbine blade simulation components, blade static strength assessment, and fracture speed prediction, thus filling gaps in related research in this field. Attached Figure Description

[0030] Figure 1 This is a flowchart of the present invention;

[0031] Figure 2 This is a schematic diagram of the variable cross-section profile in Example 2;

[0032] Figure 3 This is a schematic diagram of the shape of the simulated component in Example 2;

[0033] Figure 4 This is a schematic diagram of the simulated load boundary in Example 2;

[0034] Figure 5 This is a schematic diagram of the simulation results in Example 2. Detailed Implementation

[0035] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0036] It should be noted that various aspects of embodiments within the scope of the appended claims are described below. It will be apparent that the aspects described herein can be embodied in a wide variety of forms, and any particular structure and / or function described herein is merely illustrative. Based on this disclosure, those skilled in the art will understand that one aspect described herein can be implemented independently of any other aspect, and two or more of these aspects can be combined in various ways. For example, any number of aspects set forth herein can be used to implement the device and / or practice the method. Additionally, this device and / or method can be implemented using other structures and / or functionalities besides one or more of the aspects set forth herein.

[0037] In addition, specific details are provided in the following description to facilitate a thorough understanding of the examples, and those skilled in the art can understand the specific meaning of the above terms in this application based on the specific circumstances.

[0038] Example 1

[0039] Figure 1 This paper presents a design method for aero-engine turbine blade simulation components based on full-field stress gradient. Building upon existing simulation design methods, it considers the full-field stress gradient distribution across the blade body to verify whether aero-engine blades fail at critical speeds. Simultaneously, it introduces a safety reserve factor, defines a critical section, and extracts the temperature and stress distribution under the critical section as a special critical surface for the simulation component's evaluation. By reflecting the geometric characteristics of single-crystal turbine blades and ensuring consistency between the simulated critical section and the characteristic stress values ​​of the real blade's critical section, this method can serve and support the design of aero-engine turbine blade simulation components, blade static strength assessment, and fracture speed prediction, thus addressing shortcomings in this field. The specific steps include:

[0040] S1: Obtain the geometric model, operating conditions, and material parameters of the turbine blade;

[0041] Among them, the geometric model refers to the three-dimensional geometric model of the turbine blade designed according to actual engineering requirements with the help of computer-aided design software.

[0042] The operating conditions include the rotational speed and temperature field of the turbine blades under actual working conditions.

[0043] In this embodiment, the operating speed of the turbine blades is obtained through statistical analysis of the operating data of the external engine, and the operating temperature field of the turbine blades is obtained through heat transfer analysis or actual measurement.

[0044] The material parameters include the density of the turbine blade material, the elastic modulus at the operating temperature, and Poisson's ratio.

[0045] In this embodiment, the material parameters are obtained through material performance tests or material data sheets.

[0046] S2: Based on the geometric model, operating conditions, and material parameters of the turbine blade, establish a finite element analysis model of the turbine blade, calculate the safety reserve coefficient of each section, and determine the critical section; obtain the elastic modulus and Poisson's ratio data of the critical part, as well as the normal stress distribution of the critical section, and calculate the normal tensile force F borne by the critical section;

[0047] The safety reserve factor refers to the ratio of the yield strength of the material in the

[001] direction at that temperature to the simulated normal stress, and its calculation expression is as follows:

[0048]

[0049] In the formula, n is the safety reserve coefficient, σ 0.2 S represents the yield strength of the material in the

[001] direction. 11 The normal stress obtained by simulation is consistent with the direction of the material

[001] .

[0050] The method for determining the critical section specifically includes the following steps:

[0051] S21: Based on the finite element analysis model of the turbine blade, the stress and temperature distribution of each section of the turbine blade are obtained, and the safety reserve factor of each section is calculated.

[0052] S22: Based on the simulation results and the calculated safety reserve coefficient, the danger point is obtained, and the cross section where the danger point is located is determined as the dangerous section of the blade.

[0053] The formula for calculating the normal tensile force F borne by the critical section is as follows:

[0054]

[0055] In the formula, σ is the average stress at the critical section; s is the area of ​​the critical section.

[0056] S3: Based on the determined location of the critical section, extract the temperature of the critical section and the safety reserve factor of the section, and perform equivalent stress on different sections of the blade at the temperature under the critical section to obtain a three-dimensional model of the simulated test section.

[0057] S31: Based on the same safety reserve factor and temperature, and combined with the stress gradient distribution of the blade, the material strength change caused by different blade temperature distribution is used as the basis for design consideration of the simulation part. The characteristic stress of the three-dimensional model of the simulation part test section is determined, and then the geometric dimensions of the three-dimensional model of the simulation part test section are determined.

[0058] S32: Perform finite element analysis on the simulated test section to determine the shape of the three-dimensional model of the simulated test section;

[0059] S33: Lofting out each section of the simulated part to obtain a three-dimensional model of the test section of the simulated part;

[0060] The formulas for calculating the area of ​​other cross sections are as follows:

[0061]

[0062] In the formula, s' is the area of ​​other cross sections, n′ is the safety reserve factor of other cross sections, n0 is the safety reserve factor of the dangerous cross section, and s is the area of ​​the dangerous cross section.

[0063] In this embodiment, the material parameters of the simulation component are consistent with the material parameters corresponding to the obtained critical point temperature.

[0064] S4: Based on the three-dimensional model of the simulated test section obtained by lofting, clamping ends that conform to the tensile test specimen standard are designed at both ends of the simulated part, and it is ensured that the characteristic stress of the simulated part's critical section is consistent with that of the turbine blade's critical section under the action of normal tensile force F at the critical section.

[0065] Example 2

[0066] In this embodiment, based on the analysis of the takeoff state of a certain type of aero-engine turbine blade, a design method for a turbine blade tensile simulation component considering blade stress gradient is provided, specifically including the following steps:

[0067] T1: Obtain the geometric model, operating conditions, and material parameters of the turbine blade;

[0068] In this embodiment, the temperature range of the turbine blade's takeoff temperature field is 875℃~1171℃, and the rotational speeds corresponding to the two temperature values ​​are 48150r / min and 63100r / min, respectively. The simulation results of the temperature and stress fields of the blade's geometric model are shown in Table 1. The blade material is DD6 single-crystal nickel-based superalloy, and based on the "China Aviation Materials Handbook", its density is 8780kg / m³. 3 Tables 2-4 show the elastic modulus, Poisson's ratio, and yield strength of single-crystal nickel-based superalloys as a function of temperature.

[0069] Table 1 Temperature and stress fields of the blade cross section

[0070]

[0071]

[0072] Table 2 Elastic modulus of single-crystal nickel-based superalloys at different temperatures and orientations

[0073]

[0074] Table 3 Poisson's ratios of single-crystal nickel-based superalloys at different temperatures and orientations.

[0075]

[0076] Table 4 Tensile strength of single-crystal nickel-based superalloys at different temperatures and orientations

[0077]

[0078] T2: Based on the geometric model, operating conditions, and material parameters of the turbine blade, a finite element analysis model of the turbine blade is established, the safety reserve coefficient of each section is calculated, and the critical section is determined; the elastic modulus and Poisson's ratio data of the critical part, as well as the normal stress distribution of the critical section, are obtained, and the normal tensile force F borne by the critical section is calculated;

[0079] T21: Based on the finite element analysis model of turbine blades, the stress and temperature distribution of each section of the turbine blade are obtained, and the safety reserve factor of each section is calculated.

[0080] T22: Based on the simulation results and the calculated safety reserve coefficient, the danger point is obtained, and the cross section where the danger point is located is determined as the dangerous section of the blade.

[0081] In this embodiment, the blade reserve coefficient values ​​at different cross-sectional heights are obtained through the safety reserve coefficient, as shown in Table 5.

[0082] Table 5 Risk Factors for Different Blade Section Heights

[0083]

[0084]

[0085] According to Table 5, the safety reserve factor is the smallest when the blade section height is 84.5 mm, and therefore its section is determined to be the dangerous section.

[0086] T3: Based on the determined location of the critical section, the temperature of the critical section and the safety reserve factor of the section are extracted. The characteristic stresses of different sections of the blade are equivalent to the temperature under the critical section to obtain the three-dimensional model of the test section of the simulated part.

[0087] Based on the determined critical section, its cross-sectional area is 44.02 mm². 2 At 48150 r / min, the average stress of the section is 291 MPa, and the average temperature is 1098℃. The critical section is subjected to normal tensile force. F represents the load borne by that section.

[0088] T31: Based on the same safety reserve factor and temperature, combined with the stress gradient distribution of the entire blade, the material strength change caused by different blade temperature distribution is used as the basis for design consideration of the simulation part, the characteristic stress of the three-dimensional model of the simulation part test section is determined, and then the geometric dimensions of the three-dimensional model of the simulation part test section are determined.

[0089] T32: Perform finite element analysis on the simulated component test section to determine the shape of the three-dimensional model of the simulated component test section;

[0090] T33: Lofting out each section of the simulated part to obtain a three-dimensional model of the test section of the simulated part;

[0091] In this embodiment, a cylindrical variable cross-section design is used to simulate the component, with a minimum cross-sectional area of ​​44.02 mm². 2 The radius is 3.7432 mm, and the temperature of 1098℃ is selected as the constant design temperature for the simulation component. This ensures that the characteristic stress of the cross-section of the variable cross-section simulation component is consistent with that of the blade cross-section, while also ensuring that the reserve coefficient is consistent. Table 6 shows the calculation results of the cylindrical radius of the blade cross-section under strain.

[0092] Table 6 Variable Cross-Section Radius of Simulated Parts

[0093]

[0094]

[0095] When the cross-sectional height is greater than 98.5 mm, the radius of the strain-resistance section cylinder is too large. Therefore, data with a cross-sectional height greater than 98.5 mm are not considered when designing the simulation part.

[0096] Figure 2 The diagram shown is a schematic of the variable cross-section profile.

[0097] T4: Based on the three-dimensional model of the simulated test section obtained by lofting, clamping ends that conform to the tensile test specimen standard are designed at both ends of the simulated part, and it is ensured that the characteristic stress of the simulated part's critical section is consistent with that of the turbine blade's critical section under the action of normal tensile force F at the critical section, thus completing the design of the simulated part.

[0098] In this embodiment, the clamping end has a radius of 7mm and a length of 10mm, and a transition section is provided between it and the variable cross-section. The shape of the simulated part is as follows. Figure 3 As shown, its processing technology is the same as that of a real turbine blade, and its microstructure is the same as that of the test section.

[0099] Based on the design simulation, simulation was performed on the critical section under a normal tensile force F. The results showed that the

[001] direction of the DD6 single-crystal nickel-based superalloy was aligned with the axial direction, and the normal tensile force F on the critical section was 12809.82 N at a rotational speed of 48150 r / min. When the simulation temperature was the average temperature of the critical section (1098℃), the elastic modulus of the DD6 single-crystal nickel-based superalloy in the

[001] direction was 82.098 GPa, and the Poisson's ratio was 0.413. The load boundary conditions were as follows: Figure 4 As shown, the simulation results are as follows: Figure 5 As shown.

[0100] In the description of this specification, references to terms such as "one embodiment," "some embodiments," "example," "specific example," or "some examples," etc., indicate that a specific feature, structure, material, or characteristic described in connection with that embodiment or example is included in at least one embodiment or example of the invention. In this specification, the illustrative expressions of the above terms do not necessarily refer to the same embodiment or example. Furthermore, the specific features, structures, materials, or characteristics described may be combined in any suitable manner in one or more embodiments or examples.

[0101] Although embodiments of the present invention have been shown and described above, it is understood that the above embodiments are exemplary and should not be construed as limiting the present invention. Those skilled in the art can make changes, modifications, substitutions and variations to the above embodiments within the scope of the present invention.

Claims

1. A design method for a dedicated turbine blade simulator based on full-field stress gradient, characterized in that, Includes the following steps: S1: Obtain the geometric model, operating conditions, and material parameters of the turbine blade; S2: Based on the geometric model, operating conditions, and material parameters of the turbine blade, establish a finite element analysis model of the turbine blade, calculate the safety reserve coefficient of each section, and determine the critical section; obtain the elastic modulus and Poisson's ratio data of the critical part, as well as the normal stress distribution of the critical section, and calculate the normal tensile force F borne by the critical section; S3: Based on the determined location of the critical section, extract the temperature of the critical section and the safety reserve factor of the section, and perform equivalent stress on different sections of the blade at the temperature under the critical section to obtain a three-dimensional model of the simulated test section. S31: Based on the same safety reserve factor and temperature, and combined with the stress gradient distribution of the blade, the material strength change caused by different blade temperature distribution is used as the basis for design consideration of the simulation part. The characteristic stress of the three-dimensional model of the simulation part test section is determined, and then the geometric dimensions of the three-dimensional model of the simulation part test section are determined. S32: Perform finite element analysis on the simulated test section to determine the shape of the three-dimensional model of the simulated test section; S33: Lofting out each section of the simulated part to obtain a three-dimensional model of the test section of the simulated part; S4: Based on the three-dimensional model of the simulated test section obtained by the layout, clamping ends that conform to the tensile test specimen standard are designed at both ends of the simulated part, and it is ensured that the characteristic stress of the simulated part's critical section is consistent with that of the turbine blade's critical section under the action of the normal tensile force F at the critical section. In S2, the calculation expression for the normal tensile force F borne by the critical section is: In the formula, This represents the average stress at the critical section. The critical cross-sectional area; In S3, the calculation expression for the area of ​​other cross-sections is as follows: In the formula, For other cross-sectional areas, For the safety reserve factor of other cross sections, The safety reserve factor for the dangerous section. This represents the critical cross-sectional area.

2. The design method for OEI-specific turbine blade simulation components based on full-field stress gradient according to claim 1, characterized in that, In S1, the geometric model refers to a three-dimensional geometric model of a turbine blade designed using computer-aided design software according to actual engineering requirements.

3. The design method for OEI-specific turbine blade simulation components based on full-field stress gradient according to claim 1, characterized in that, In S1, the operating conditions include the rotational speed and temperature field of the turbine blade under actual working conditions; wherein, the operating rotational speed of the turbine blade is obtained through statistical analysis of the operating data of the external engine, and the operating temperature field of the turbine blade is obtained through heat transfer analysis or actual measurement.

4. The design method for OEI-specific turbine blade simulation components based on full-field stress gradient according to claim 1, characterized in that, In S1, the material parameters include the density of the turbine blade material, the elastic modulus at the operating temperature, and the Poisson's ratio; wherein, the material parameters are obtained through material performance tests or material data sheets.

5. The design method for an OEI-specific turbine blade simulation component based on full-field stress gradient according to claim 1, characterized in that, In S2, the safety reserve factor refers to the ratio of the yield strength of the material in the [001] direction at that temperature to the simulated normal stress, and its calculation expression is: In the formula, As a safety reserve factor, The yield strength of the material in the [001] direction. The normal stress is obtained through simulation to align with the material orientation.

6. The design method for an OEI-specific turbine blade simulation component based on full-field stress gradient according to claim 1, characterized in that, In step S2, the method for determining the critical section specifically includes the following steps: S21: Based on the finite element analysis model of the turbine blade, the stress and temperature distribution of each section of the turbine blade blade in the whole field are obtained, and the safety reserve factor of each section is calculated. S22: Based on the simulation results and the calculated safety reserve coefficient, the danger point is obtained, and the cross section where the danger point is located is determined as the dangerous section of the blade.

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