A stress-aerodynamic integrated design method for a gas turbine axial compressor blade

Abstract

The purpose of this invention is to provide a stress-aerodynamic fusion design method for axial compressor blades of gas turbines, belonging to the field of gas turbines. This invention comprehensively considers the influence of compressor blade design on aerodynamic performance and blade stress distribution. Through experimental verification, it effectively solves the challenge of balancing compressor performance and reliability. A relationship model between compressor blade design parameters and blade stress distribution is established, realizing the organic integration of blade stress distribution optimization and compressor aerodynamic design, effectively improving the precision of compressor design and shortening the design cycle. Furthermore, this invention is not limited to gas turbine compressors but is also applicable to the design process of various industrial axial compressors and aero-engine axial compressors.

Description

Technical Field

[0001] The present invention relates to a gas turbine design method, specifically a moving blade design method. Background Technology

[0002] As one of the three core components of a gas turbine, the performance and reliability of the compressor directly determine the overall performance level of the gas turbine. With the continuous development of gas turbine technology, the requirements for the aerodynamic performance and load levels of compressor components are also rapidly increasing. To meet these performance requirements, the three-dimensional configuration of compressor blades is becoming increasingly complex, leading to increasingly severe reliability issues. This is particularly true for marine and industrial gas turbines, which have long service lives and high reliability requirements, posing a significant challenge to the design of high-performance compressor blades. Furthermore, the gas in the compressor flows at high speed under strong adverse pressure gradients, exhibiting strong three-dimensional and nonlinear characteristics in its internal flow field. Any adjustment to the blade structure can cause flow separation within the compressor, leading to performance degradation and affecting the unit's operating efficiency and stability. Achieving a balance between compressor aerodynamic performance and blade reliability has become a core challenge in the design of modern advanced gas turbine compressors. Therefore, it is necessary to explore new methods and technologies that can simultaneously consider the stress level and aerodynamic loss level of compressor blades, enabling the compressor to achieve high reliability while meeting aerodynamic performance requirements. Summary of the Invention

[0003] The purpose of this invention is to provide a gas turbine axial compressor blade stress-aerodynamic fusion design method that can solve the problem of balancing the aerodynamic performance of gas turbine compressors and the reliability of blades.

[0004] The objective of this invention is achieved as follows: This invention discloses a stress-aerodynamic fusion design method for axial compressor blades of gas turbines, characterized by comprising the following steps: (1) Initial blade design: Based on the compressor flow design results, the blade geometry is designed, and the two-dimensional blade profiles of each characteristic section along the blade height are designed. The centroid position of each section blade profile is calculated, and the three-dimensional blade design is completed by the centroid stacking method to obtain the initial aerodynamic blade before blade stress control. (2) Three-dimensional CFD calculation and analysis: Perform three-dimensional CFD calculation of the compressor flow to obtain the overall aerodynamic performance parameters of the compressor and the pressure and temperature distribution of each row of blades. (3) Static analysis of blades: For the compressor blades, centrifugal load, aerodynamic load and temperature load are extracted from the design conditions selected from the three-dimensional CFD calculation results, and static finite element calculation is performed to obtain the stress distribution of the blades. (4) Determine whether the overall aerodynamic performance parameters of the compressor meet the aerodynamic design requirements based on the three-dimensional CFD calculation results, and extract the maximum stress value on the blade head based on the static finite element calculation results of the moving blade. Maximum stress value on the back of the blade Determine whether the uniformity of blade stress meets the convergence condition of blade stress design; if it does, the scheme is the final scheme; if it does not, continue the optimization design according to step (5); (5) Blade shape optimization design: including two-dimensional blade shape optimization along each characteristic section of the blade height and blade stacking line optimization. By optimizing the two-dimensional blade shape of each characteristic section, the position of maximum stress on the blade is adjusted to the ideal position. By optimizing the stacking line, the uniformity of stress on the blade is adjusted to obtain the optimized blade scheme. Steps (2), (3), and (4) are repeated. After repeated iterations, the final blade scheme that meets the design requirements is obtained.

[0005] The present invention may also include: 1. The calculation method for determining whether the blade stress uniformity meets the blade stress design convergence condition in step (4) is as follows: Blade stress uniformity is measured by the stress deviation between the two sides of the blade. calculate: ; Blade stress design convergence conditions: ; in Value Or adjust according to the actual situation of the compressor.

[0006] 2. In step (5), the optimization of the two-dimensional airfoil of each characteristic section adopts the geometric shape parameters related to the airfoil thickness as optimization parameters, including chord length, maximum thickness position, relative maximum thickness, leading edge radius, trailing edge radius. When adjusting the optimization parameters, the angle of attack and the lag angle of the optimized two-dimensional airfoil are calculated simultaneously to adjust its inlet geometric angle, outlet geometric angle, airfoil bend angle, and airfoil aerodynamic shape parameters, so as to ensure that the aerodynamic performance of the two-dimensional airfoil of this section is minimally affected.

[0007] 3. In step (5), adjust the position of maximum stress on the blade to the ideal position by pointing to the maximum stress value on the blade base. Maximum stress value on the back of the blade They are all located in the leaf root area on both sides of the leaf base and leaf back, and are not close to the leading edge and trailing edge.

[0008] 4. The optimization of the stacking line mentioned in step (5) is carried out using the following method: (5.1) Fit the accumulation line to a straight line passing through the centroid of the blade root and blade tip section. Take the centroid of the blade root section as the base point and rotate this straight line along the circumferential direction of the compressor by the following angle. , The positive direction is towards the back of the leaf; (5.2) The value is selected as follows: During the first optimization, assign A constant value , The value is preliminarily estimated based on experience, taking into account the stress deviation on both sides of the blade in the initial design before optimization; During the second optimization... The initial design of the blade and the stress deviation values ​​on both sides of the blade body. And the stress deviation values ​​on both sides of the blade in the first optimization scheme. Perform linear interpolation to obtain The calculation formula, and the calculation Perform a second optimization, such as the stress deviation value on both sides of the blade after the second optimization. If the calculation results fail to meet the convergence condition for blade stress design, then ( ,0), ( , ), ( , Perform a second interpolation to obtain a new... Calculation formula, and calculation Then, a third optimization was performed. (5.3) If the requirements are still not met after three optimizations, then adjust... Repeat step (5.2) for the value.

[0009] The advantages of this invention are: 1. This invention comprehensively considers the influence of compressor blade design on aerodynamic performance and blade stress distribution. Through experimental verification, it effectively solves the problem of balancing compressor performance and reliability.

[0010] 2. This invention establishes a relationship model between the compressor blade styling parameters and the blade stress distribution, enabling the blade stress distribution to be quickly adjusted in the blade styling design process. This achieves the organic integration of blade stress distribution optimization and compressor aerodynamic design, effectively improving the precision of compressor design and shortening the design cycle.

[0011] 3. This invention is not limited to gas turbine compressors, but is also applicable to the design process of various industrial axial compressors and aero-engine axial compressors. Attached Figure Description

[0012] Figure 1 This is a flowchart of the present invention. Detailed Implementation

[0013] The invention will now be described in more detail with reference to the accompanying drawings: Combination Figure 1 The present invention is achieved through the following steps: Step 1: Initial blade design. Based on the compressor flow path design results, the blade geometry is designed, and two-dimensional airfoils are designed for each characteristic section along the blade height. The centroid position of each airfoil section is calculated, and the three-dimensional blade design is completed by the centroid stacking method to obtain the initial aerodynamic blade design before blade stress control. Step 2: 3D CFD Calculation and Analysis. Perform a 3D CFD calculation of the compressor flow path to obtain the overall aerodynamic performance parameters of the compressor and the pressure and temperature distribution of each blade profile. Step 3: Blade Static Analysis. For the compressor blade, centrifugal load, aerodynamic load, and temperature load are extracted from the design conditions selected in the 3D CFD calculation results, and static finite element analysis is performed to obtain the stress distribution of the moving blade; Step 4: Determine if design requirements are met. First, based on the 3D CFD calculation results, determine whether the overall aerodynamic performance parameters of the compressor meet the aerodynamic design requirements. Then, extract the maximum stress value on the blade head based on the static finite element calculation results of the moving blades. Maximum stress value on the back of the blade To determine whether the uniformity of blade stress meets the design convergence condition for blade stress.

[0014] Blade stress uniformity is measured by the stress deviation between the two sides of the blade. calculate: ; Blade stress design convergence conditions: in Common values It can also be adjusted according to the actual situation of the compressor.

[0015] If the conditions are met, the solution is the final solution; if not, proceed to step five for further optimization. Step 5: Blade Shape Optimization Design. This includes 2D blade shape optimization along various characteristic sections of the blade height and blade overlap line optimization. Specifically: Optimizing two-dimensional airfoils at various characteristic sections requires considering both aerodynamic and structural effects. Typically, geometric parameters related to airfoil thickness are used as optimization parameters, such as chord length, maximum thickness location, relative maximum thickness, leading edge radius, and trailing edge radius. When adjusting these optimization parameters, the angle of attack and lag angle of the optimized two-dimensional airfoil must be calculated simultaneously. This information is used to adjust aerodynamic parameters such as the inlet geometry, outlet geometry, and airfoil bend angle, ensuring minimal impact on the aerodynamic performance of the airfoil at that cross-section. Optimizing the two-dimensional airfoil at various characteristic sections adjusts the location of maximum stress in the blade to the ideal position, at which point the maximum stress value on the moving blade tip is determined. Maximum stress value on the back of the blade They are all located in the leaf root area on both sides of the leaf base and leaf back, and are not close to the leading edge and trailing edge.

[0016] The following method is used to optimize the accumulation line: (1) Fit the accumulation line to a straight line passing through the centroid of the blade root and blade tip section. Take the centroid of the blade root section as the base point and rotate the straight line along the circumferential direction of the compressor by the following angle. , The direction of rotation towards the back of the leaf is the positive direction; (2) The value is selected as follows: During the first optimization, assign A constant value , The value can be preliminarily estimated based on experience according to the stress deviation on both sides of the blade in the initial scheme before optimization; the first During the second optimization... The initial design of the blade ( Stress deviation value on both sides of the blade And the blades of the first optimization scheme ( Stress deviation value on both sides of the blade Perform linear interpolation to obtain The calculation formula, and the calculation A second optimization is performed, such as the stress deviation value on both sides of the blade after the second optimization. If the calculation results fail to meet the convergence condition for blade stress design, then ( ,0), ( , ), ( , Perform a second interpolation to obtain a new... Calculation formula, and calculation (3) Usually, the blade stress design convergence condition can be met after three optimizations. If it still cannot be met, it needs to be adjusted. Repeat step (2) until the blade stress uniformity meets the requirements.

[0017] Through the aforementioned blade shape optimization design work, an optimized blade scheme is obtained, and steps two, three, and four are repeated. Through repeated iterations of these steps, the final blade scheme that meets the design requirements is obtained.

[0018] The stress-aerodynamic fusion design method for gas turbine axial compressor blades proposed in this invention is universal and is not limited to gas turbine compressors. It is also applicable to the design process of various industrial axial compressors and aero-engine axial compressors.

Claims

1. A method for stress-aerodynamic fusion design of axial compressor blades for gas turbines, characterized by: Includes the following steps: (1) Initial blade design: Based on the compressor flow design results, the blade geometry is designed, and the two-dimensional blade profiles of each characteristic section along the blade height are designed. The centroid position of each section blade profile is calculated, and the three-dimensional blade design is completed by the centroid stacking method to obtain the initial aerodynamic blade before blade stress control. (2) Three-dimensional CFD calculation and analysis: Perform three-dimensional CFD calculation of the compressor flow to obtain the overall aerodynamic performance parameters of the compressor and the pressure and temperature distribution of each row of blades. (3) Static analysis of blades: For the compressor blades, centrifugal load, aerodynamic load and temperature load are extracted from the design conditions selected from the three-dimensional CFD calculation results, and static finite element calculation is performed to obtain the stress distribution of the blades. (4) Determine whether the overall aerodynamic performance parameters of the compressor meet the aerodynamic design requirements based on the three-dimensional CFD calculation results, and extract the maximum stress value on the blade head based on the static finite element calculation results of the moving blade. Maximum stress value on the back of the blade Determine whether the uniformity of blade stress meets the convergence condition of blade stress design; if it does, the scheme is the final scheme; if it does not, continue the optimization design according to step (5); (5) Blade shape optimization design: including two-dimensional blade shape optimization along each characteristic section of the blade height and blade stacking line optimization. By optimizing the two-dimensional blade shape of each characteristic section, the position of maximum stress on the blade is adjusted to the ideal position. By optimizing the stacking line, the uniformity of stress on the blade is adjusted to obtain the optimized blade scheme. Steps (2), (3), and (4) are repeated. After repeated iterations, the final blade scheme that meets the design requirements is obtained.

2. The gas turbine axial compressor blade stress-aerodynamic fusion design method according to claim 1, characterized in that: The calculation method for determining whether the blade stress uniformity meets the blade stress design convergence condition in step (4) is as follows: Blade stress uniformity is measured by the stress deviation between the two sides of the blade. calculate: ; Blade stress design convergence condition: ; wherein values or adjusted according to the actual situation of the compressor.

3. A stress-aerodynamic blended design method for a gas turbine axial compressor rotor blade airfoil as recited in claim 1, characterized by: In step (5), the optimization of the two-dimensional airfoil of each characteristic section adopts geometric shape parameters related to the airfoil thickness as optimization parameters, including chord length, maximum thickness position, relative maximum thickness, leading edge radius, trailing edge radius. When adjusting the optimization parameters, the angle of attack and the lag angle of the optimized two-dimensional airfoil are calculated simultaneously to adjust its inlet geometric angle, outlet geometric angle, airfoil bend angle, and airfoil aerodynamic shape parameters, so as to ensure that the aerodynamic performance of the two-dimensional airfoil of this section is minimally affected.

4. A stress-aerodynamic blended design method for a gas turbine axial compressor rotor blade airfoil as defined in Claim 1, characterized by: The adjustment of the maximum stress position of the blade body to the ideal position in step (5) means that the maximum stress value on the blade body of the blade and the maximum stress value on the blade back are both located in the blade root area on both sides of the blade pad and the blade back, and are not close to the leading edge and the trailing edge.

5. A stress-aerodynamic blended design method for a gas turbine axial compressor rotor blade airfoil as defined in Claim 1, characterized by: The optimization of the stacking line in step (5) is achieved using the following method: (5.1) Fit the accumulation line to a straight line passing through the centroid of the blade root and blade tip section. Take the centroid of the blade root section as the base point and rotate this straight line along the circumferential direction of the compressor by the following angle. , The positive direction is towards the back of the leaf; (5.2) The value is selected as follows: During the first optimization, assign A constant value , The value is preliminarily estimated based on experience, taking into account the stress deviation on both sides of the blade in the initial design before optimization; the first During the second optimization... The initial design of the blade and the stress deviation values ​​on both sides of the blade body. And the stress deviation values ​​on both sides of the blade in the first optimization scheme. Perform linear interpolation to obtain The calculation formula, and the calculation Perform a second optimization, such as the stress deviation value on both sides of the blade after the second optimization. If the calculation results fail to meet the convergence condition for blade stress design, then ( ,0), ( , ), ( , Perform a second interpolation to obtain a new... Calculation formula, and calculation Then, a third optimization was performed. (5.3) If the condition is not satisfied after 3 optimizations, adjust the value of and go back to step (5.2).

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