Gas turbine blade cover quantity optimization method comprehensively considering blade strength and aerodynamic performance

Abstract

The invention discloses a gas turbine blade cover amount optimization method comprehensively considering blade strength and aerodynamic performance. Blade cover amount adjustment and cold and hot state conversion are taken as a complete optimization process; wherein the cover amount of the blade is an optimization variable, and adjustment is carried out according to section deformation obtained through blade numerical analysis; the blade profile after cover amount adjustment is a first round optimization target blade profile; the maximum stress difference between a blade basin and a blade back in a target blade profile numerical analysis result and a pneumatic design index are cover quantity optimization targets; a deformed blade profile obtained after a thermal load is applied to the target blade profile in the second wheel casing quantity optimization is an initial thermal state blade profile, and a difference value between the thermal state blade profile and the target blade profile is a cold and hot state conversion optimization target; the maximum stress or deformation under the thermal-state blade profile serves as a final optimization target, and the optimized thermal-state blade profile is obtained. By means of the method, the cold-state blade profile and the hot-state blade profile can be optimized at the same time, and the problem that the strength and aerodynamic performance of the blade profile are reduced after a traditional cover amount adjustment and cold and hot state conversion process is carried out is solved.

Description

Technical Field

[0001] The present invention belongs to the technical field of gas turbine structural strength and aerodynamics, and particularly relates to a method for optimizing the shroud amount of blades by comprehensively considering blade strength and aerodynamic performance. Background Art

[0002] The traditional process for optimizing the shroud amount of gas turbine rotor blades is as follows: First, the shroud amount of the initial blade profile is adjusted to ensure that the strength reserve of the blade meets the requirements; then, the cold-state airfoil is obtained through cold-hot state conversion to ensure the aerodynamic performance of the blade. The commonly used method for adjusting the shroud amount of gas turbine blades is to keep the profile cross-sectional shape unchanged along the blade height direction and adjust the displacements (shroud amounts) of each cross-section along the axial and circumferential directions, thereby reducing the maximum stress or deformation on the blade and ensuring uniform stress distribution on the blade basin and blade back; the displacement amounts of each airfoil cross-section along the axial and circumferential directions are linearly distributed along the blade height, with the maximum at the blade tip and zero at the blade root cross-section. The traditional cold-hot state conversion methods are divided into two types. One starts from the design of the hot-state airfoil, and by comparing with the target airfoil, the airfoil parameters are optimized to ensure the initial aerodynamic performance indicators. However, considering the cold-hot state conversion of the blade as a completely linear process and ignoring the large deformations caused by geometric non-linearity often results in errors in the results. The other starts from the design of the cold-state airfoil, and by optimizing the airfoil parameters and their strength performance indicators (such as maximum stress, deformation, etc.), the strength performance of the blade is ensured. However, since the compensation effect of the centrifugal load on the aerodynamic moment is not considered separately, it may lead to a decrease in aerodynamic performance. At the same time, the complex iterative process will also increase the error of the results.

[0003] Therefore, to improve the defects brought by the above traditional methods, a method for optimizing the shroud amount by comprehensively considering blade strength and aerodynamic performance is needed. Summary of the Invention

[0004] The purpose of the present invention is to provide a method for optimizing the shroud amount of gas turbine blades by comprehensively considering blade strength and aerodynamic performance, so as to avoid the disadvantages of the corresponding hot-state airfoil strength performance degradation or the reduction of the aerodynamic performance (such as aerodynamic efficiency, flutter margin, etc.) of the pre-deformed cold-state airfoil after the traditional shroud amount adjustment and cold-hot state conversion processes.

[0005] To achieve the above purpose, the present invention adopts the following technical solutions: A method for optimizing the shroud amount of gas turbine blades by comprehensively considering blade strength and aerodynamic performance, comprising the following steps: (1) Obtain the initial cold-state airfoil of the aerodynamic design, and use simulation software to perform numerical analysis on its response under structural and aerodynamic loads to obtain the axial and circumferential deformation amounts of each airfoil cross-section, and obtain the cold-state working airfoil; (2) Modify the cold-state working airfoil according to the axial and circumferential deformation amounts of each cross-section in step (1), and use modeling software to convert the modified airfoil data into a three-dimensional model to obtain the target airfoil; (3) Use simulation software to perform numerical analysis on the target airfoil obtained in step (2) to obtain the difference in the maximum stresses between the blade suction side and pressure side; (4) Determine whether the difference in the maximum stresses between the blade suction side and pressure side obtained in step (3) meets the requirements. If it meets, proceed to the next step; otherwise, repeat step (2) according to the result; (5) Check whether the aerodynamic performance of the target airfoil meets the design specifications. If it meets, proceed to the next step; otherwise, repeat step (2) according to the result; (6) Apply a thermal load to the target airfoil obtained in step (5) and perform numerical analysis to obtain strength parameters such as the stress and deformation of the initial hot-state airfoil; (7) If the maximum deviation of each cross-section of the airfoil obtained in step (6) from the target airfoil meets the requirements, proceed to the next step; otherwise, adjust the axial and circumferential deformation amounts of each cross-section, and use this airfoil as the input to repeat step (6); (8) If the strength parameters such as the stress and deformation obtained in step (6) meet the requirements, then this airfoil is the optimized hot-state airfoil, proceed to the next step; otherwise, adjust the axial and circumferential deformation amounts of each cross-section, and use this airfoil as the input to repeat step (6); (9) According to the optimized hot-state airfoil obtained in step (7) and its shrouding value, inversely deduce the optimized cold-state airfoil.

[0006] In the said step (1), the axial and circumferential deformation amounts of each cross-section of the cold-state working airfoil are the average values in the cylindrical coordinate system.

[0007] In the said step (2), add the average deformation values obtained in step (1) to the axial and circumferential coordinates of the initial cold-state airfoil data, while keeping its radial coordinate and airfoil cross-section line form unchanged.

[0008] In the said step (3), the deformation values of each cross-section of the target airfoil include the average values in the cylindrical coordinate system.

[0009] In the said step (4), the maximum stresses of the blade suction side and pressure side of the target airfoil are respectively the maximum stress values of the elements in the suction side and pressure side regions extracted by the simulation software.

[0010] In the said step (5), the aerodynamic performance index of the target airfoil is extracted by the simulation software and compared with the design values (such as aerodynamic efficiency, flutter margin, etc.).

[0011] In step (6), without considering the radial deformation difference of the same cross-section, the difference between the initial hot-state airfoil and the target airfoil is used as the coordinate data of the initial hot-state airfoil minus the coordinate data of the target airfoil, and the sum of its absolute values is used as the basis for the next judgment.

[0012] In step (7), when the difference between the initial hot-state airfoil and the target airfoil is less than a given value (e.g., 0.02), it is considered that the shroud amount optimization of the initial hot-state airfoil is completed. Among them, when the sum of the absolute values of the current airfoil difference is greater than the value of the previous round, the sign of the airfoil coordinate data difference obtained in step (7) is changed and used as the input of step (6).

[0013] In step (8), when the maximum stress or deformation of the initial hot-state airfoil is less than a given value (e.g., 600 Mpa), it is considered that the strength index of this target airfoil meets the requirements, and the airfoil coordinate data and airfoil difference obtained in step (8) are recorded and used as the input of step (9). Otherwise, it is used as the input of step (6).

[0014] Beneficial effects: The present invention takes the blade shroud amount adjustment and the cold-hot state conversion as a complete optimization process. Among them, the shroud amount of the blade is the optimization variable, which is adjusted according to the cross-section deformation obtained from the numerical analysis of the blade; the airfoil that meets the strength and aerodynamic performance indicators after the shroud amount adjustment is the target airfoil; the difference between the maximum stresses on the suction side and the pressure side of the target airfoil in the numerical analysis result and the airfoil difference from the original cold-state airfoil are the shroud amount optimization targets; the target airfoil is corrected according to the data point displacement in the numerical analysis result after applying the thermal load to obtain the initial hot-state airfoil; the difference between the initial hot-state airfoil and the target airfoil is the cold-hot state shroud amount optimization target; the maximum stress or deformation under the hot-state airfoil is the final optimization target. Through the above method, it is possible to avoid the problem of the decline of the strength and aerodynamic performance indicators of the optimized cold-state and hot-state airfoils after passing through the traditional shroud amount adjustment and cold-hot state conversion processes. Description of the Drawings

[0015] Figure 1 is the flowchart of the present invention; Figure 2 is a schematic diagram comparing the airfoils before and after the blade optimization in the embodiment of the present invention. In the figure, 1 is the initial cold-state airfoil, 2 is the target airfoil, 3 is the initial hot-state airfoil, 4 is the optimized hot-state airfoil, and 5 is the optimized cold-state airfoil. Detailed Embodiment

[0016] To make the technical solutions and their advantages of this application clearer, the following will further describe the technical solutions of this application clearly and completely in conjunction with the accompanying drawings. It can be understood that the specific embodiments described herein are only partial embodiments of this application, which are only used to explain this application and not to limit this application. It should be noted that for the convenience of description, only the parts related to this application are shown in the drawings, and other related parts can be referred to the general description. Without conflict, the embodiments in this application and the technical features in the embodiments can be combined with each other to obtain new embodiments.

[0017] In addition, unless otherwise defined, the technical terms or scientific terms used in the description of this application should be the ordinary meanings understood by those of ordinary skill in the field to which this application belongs. The similar terms such as "each" or "the" used in the description of this application should not be understood as an absolute limitation on the quantity, but should be understood as having at least one.

[0018] It should also be noted that unless otherwise clearly specified and limited, the similar terms such as "average", "correction", "conversion" used in the description of this application should be understood in a broad sense, and those skilled in the art can understand their specific meanings in this application according to the specific situation.

[0019] The following will further describe this application in detail in conjunction with the attached Figures 1 to 2 drawings.

[0020] This application relates to an optimization method for the shroud quantity of a gas turbine blade that comprehensively considers blade strength and aerodynamic performance. Its implementation process is as shown in the attached Figure 1 drawings. The steps of its application in the structural optimization of the compressor blade of a certain type of engine include: (1) For the initial airfoil of the aerodynamic design (as shown in 1 in the attached Figure 2 drawings), use the simulation software ANSYS to perform numerical analysis on its response under the design load to obtain the axial and circumferential deformation values of each airfoil section; among them, the deformation of each section of the obtained initial airfoil is the average value in the cylindrical coordinate system.

[0021] (2) According to the axial and circumferential deformation values obtained in step (1), correct the data of each section of the initial airfoil, and use the modeling software UG to convert the corrected airfoil data into a three-dimensional model to obtain the target airfoil (as shown in 2 in the attached Figure 2 drawings); among them, to ensure the convergence of the optimization target and the strength and aerodynamic performance of the airfoil, add the average deformation value obtained in step (1) to the axial and circumferential coordinates of the initial airfoil data. Since no thermal load is applied, the radial coordinate and the form of the airfoil section line remain unchanged.

[0022] (3) Use the simulation software ANSYS to perform numerical analysis on the target airfoil to obtain the difference in the maximum stress between the suction side and the pressure side regions.

[0023] (4) Judge the blade profile obtained in step (3). Use simulation software to judge whether the maximum stress difference between the blade back and the blade basin meets the set requirements. If it meets the requirements, proceed to the next step; otherwise, adjust the cross-section deformation value of the target blade profile as the input of step (2), and repeat steps (2) to (4).

[0024] (5) According to the blade profile obtained in step (3), use simulation software to judge whether the aerodynamic performance (such as aerodynamic efficiency, flutter margin, etc.) meets the set requirements. If it meets the requirements, proceed to the next step; otherwise, adjust the cross-section deformation value of the target blade profile as the input of step (2), and repeat steps (2) to (5).

[0025] (6) Use the simulation software ANSYS to apply thermal load to the target blade profile for numerical analysis. Without considering the radial displacement difference of the same cross-section, obtain the difference between each cross-section of the hot blade profile and the target blade profile, as well as the maximum deformation / stress value; among them, the difference between the hot blade profile and the target blade profile is the coordinate data of the hot blade profile at the same cross-section minus the coordinate data of the target blade profile.

[0026] (7) Judge whether the difference between the initial hot blade profile and the target blade profile in step (6) meets the set requirements. If it meets the requirements, proceed to the next step; otherwise, correct the blade profile difference obtained in step (7) as the input of step (6), and repeat steps (6) to (7); among them, when the difference between the initial hot blade profile and the target blade profile is less than the given value of 0.02, it is considered to meet the requirements. When the sum of the absolute values of the current blade profile difference is greater than the value of the previous round, change the sign of the blade profile coordinate data difference obtained in step (7) as the input of step (6).

[0027] (8) Judge whether the maximum deformation / stress in step (6) meets the requirements. If it meets the requirements, stop the optimization and obtain the optimized hot blade profile; otherwise, correct the blade profile difference obtained in step (8) as the input of step (6), and repeat steps (6) to (8); among them, the set requirements for the maximum deformation / stress can be less than a certain specific value or reach its minimum value.

[0028] (9) According to the optimized hot blade profile and the blade profile difference obtained in step (8), inversely deduce the optimized target blade profile, and then according to the blade profile difference obtained in step (5), inversely deduce the optimized cold blade profile.

Claims

1. A method for optimizing the blade cover amount of a gas turbine by comprehensively considering blade strength and aerodynamic performance, characterized in that: The following steps are involved: (1) For the initial cold blade profile of the aerodynamic design, the simulation software is used to perform numerical analysis on its response under structural and aerodynamic loads, and the axial and circumferential deformations of each blade section are obtained to obtain the cold working blade profile; (2) modifying the cold working blade profile according to the axial and circumferential deformation of each cross section in step (1) and converting the modified blade profile data into a three-dimensional model using modeling software to obtain a target blade profile; (3) Using simulation software to perform numerical analysis on the target blade profile obtained in step (2) to obtain the difference between the maximum stresses of the blade base and the blade back; (4) Determine whether the maximum stress difference between the blade base and the blade back obtained in step (3) meets the requirements. If so, proceed to the next step. Otherwise, repeat step (2) according to the result. (5) Check whether the aerodynamic performance of the target blade meets the design specifications. If so, proceed to the next step. Otherwise, repeat step (2) based on the result. (6) applying a thermal load to the target blade profile obtained in step (5) and performing a numerical analysis to obtain strength parameters such as stress and deformation of the initial hot blade profile; (7) If the maximum deviation of each section of the blade profile obtained in step (6) from the target blade profile meets the requirements, proceed to the next step. Otherwise, adjust the axial and circumferential deformation of each section and repeat step (6) with this blade profile as input. (8) If the strength parameters such as stress and deformation obtained in step (6) meet the requirements, then this blade profile is the optimized hot blade profile and the next step is performed. Otherwise, the axial and circumferential deformation of each section is adjusted and step (6) is repeated with this blade profile as input; (9) Based on the optimized hot blade profile and its shroud value obtained in step (7), the optimized cold blade profile is obtained by reverse calculation.

2. The method for optimizing the gas turbine blade cover volume by comprehensively considering blade strength and aerodynamic performance according to claim 1, characterized in that: In the step (1), the axial and circumferential deformation values ​​of each cross section of the initial blade profile are their average values ​​in the cylindrical coordinate system.

3. The method for optimizing the gas turbine blade cover volume by comprehensively considering blade strength and aerodynamic performance according to claim 1, characterized in that: In the step (2), the axial and circumferential coordinates of the initial cold blade profile data are added with the average deformation value obtained in the step (1), while the radial coordinates and the blade profile cross-sectional line form remain unchanged.

4. The method for optimizing the gas turbine blade cover volume by comprehensively considering blade strength and aerodynamic performance according to claim 1, characterized in that: In the step (3), the difference between the maximum stress of the blade base and the blade back is obtained by simulation software.

5. The method for optimizing the gas turbine blade cover volume by comprehensively considering blade strength and aerodynamic performance according to claim 1, characterized in that: In step (5), the aerodynamic performance of the blade is checked by simulation software to see whether it meets the design requirements (such as aerodynamic efficiency, flutter margin, etc.).

6. The method for optimizing the gas turbine blade cover volume by comprehensively considering blade strength and aerodynamic performance according to claim 1, characterized in that: In the step (6), the radial displacement difference of the same cross section is not considered, and the deformation value of each cross section of the target blade profile includes its average value in the cylindrical coordinate system and the individual deformation value of each data point.

7. The method for optimizing the gas turbine blade cover volume by comprehensively considering blade strength and aerodynamic performance according to claim 1, characterized in that: In the step (7), when the difference between the initial hot blade profile and the target blade profile is less than a given value (e.g., 0.02), it is considered that the requirement is met, wherein when the sum of the absolute values ​​of the current blade profile differences is greater than the value of the previous round, the blade profile coordinate data difference obtained in step (7) is changed in sign and used as the input of step (6).

8. The method for optimizing the gas turbine blade cover volume by comprehensively considering blade strength and aerodynamic performance according to claim 1, characterized in that: In the step (8), it is necessary to determine whether the maximum deformation / stress in step (6) meets the requirements. If so, the optimization is stopped to obtain the optimized hot blade profile. Otherwise, the blade profile difference obtained in step (8) is corrected and used as the input of step (6), and steps (6) to (8) are repeated; wherein the setting requirement of the maximum deformation / stress can be less than a certain specific value (for example, 600 MPa) or can reach its minimum value.

9. The method for optimizing the gas turbine blade cover volume by comprehensively considering blade strength and aerodynamic performance according to claim 1, characterized in that: In the step (9), the optimized hot blade profile and the blade profile difference obtained in step (8) are used to reversely obtain the optimized target blade profile, and then the optimized cold blade profile is reversely obtained based on the blade profile difference obtained in step (5).

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