A Blade Reverse Torsion Design Method Based on Blade Section Displacement - Characteristic Angle

Abstract

The present invention discloses a method for reverse torsion design of blades based on blade cross-section displacement - characteristic angle. The present invention comprehensively considers the influence of the aerodynamic load and centrifugal load acting on the blades during the operation of the gas turbine compressor blades on the blade deformation, extracts the coordinate points at different cross-sections of multiple blades for fitting, and converts the displacement amounts of the coordinate points at different cross-sections of the hot and cold state blades into the change of the geometric characteristic angle of the blade cross-section to realize the reverse torsion design of the cold-state machined blades, and obtain the final machining scheme of the compressor moving blades. The method proposed by the present invention can greatly improve the efficiency of blade reverse torsion design, shorten the design cycle of the gas turbine, and can also be used as an important method for the practical application of blade reverse torsion design in the fields of aviation and shipbuilding on the basis of ensuring the surface, surrounding flow conditions and overall aerodynamic performance of the compressor blades.

Description

Technical Field

[0001] The invention belongs to the field of aerodynamic design of an axial flow compressor of a gas turbine, and in particular relates to a blade anti-twist design method based on blade section displacement-characteristic angle. Background Art

[0002] Impeller machinery is developing towards the performance indicators of high load, high thrust-to-weight ratio, high efficiency and low fuel consumption. High-speed rotating blades are prone to large geometric deformation under the combined action of centrifugal load and aerodynamic load, making the instability of the flow field and aerodynamic elasticity problems more prominent, thereby shortening the axial spacing of the blades, enhancing the dynamic and static interference during the operation of multiple rows of blades, and intensifying the aerodynamic excitation of the blade surface, thereby causing high-cycle fatigue and even fatigue fracture problems. The impeller machinery blades produce geometric deformation under the influence of aerodynamic load and centrifugal load, which changes the blade deformation and causes the aerodynamic performance to shift. The blade shape is one of the most important parameters that determine the performance of the compressor. Ensuring that the blade shape is as close to the design state as possible during the operation of the compressor is an important link and guarantee for achieving its design indicators.

[0003] The deformation of compressor blades under actual working conditions will affect the blade installation angle, blade tip clearance, and inlet and outlet geometric angles, causing the flow in the compressor to deviate from the design conditions. Especially for the inlet transonic stage, blade deformation will cause the shock wave position to move, thereby deviating from its optimal design point, which will affect the compressor performance and efficiency, and even bring about problems such as aeroelastic stability. To achieve the design performance indicators of the compressor, it should be ensured that the shape of the blade is close to or consistent with the design shape after it is transformed from the cold state (processing state) to the hot state (working state). To this end, it is necessary to study the reverse process of the blade from the designed blade shape to the cold processed blade shape, which is called the reverse twist design of the blade. Therefore, providing accurate blade processing data becomes an important link in the design process to improve design efficiency and shorten the design cycle. Summary of the invention

[0004] In order to provide accurate blade processing data, the present invention proposes a blade anti-twist design method based on blade cross-section displacement-characteristic angle. This method can greatly improve the blade anti-twist design efficiency while ensuring the compressor blade surface, surrounding flow conditions and overall aerodynamic performance, thereby shortening the design cycle of the gas turbine.

[0005] To achieve the above object, the present invention provides the following solution: a blade anti-twist design method based on blade section displacement-characteristic angle, comprising the following steps:

[0006] Step 1: Perform static analysis on the theoretical blade under the aerodynamic load and centrifugal load of the design condition;

[0007] Step 2: Calculate the external force vector through a structural mechanics solver based on the pneumatic load and centrifugal load, and obtain the primary hot state deformation amount when the theoretical blade internal force vector and the external force vector are balanced;

[0008] Step 3: Extract the hot state deformation amounts of multiple cross-sections of the theoretical blade along the blade height direction, and obtain the primary hot state blade through coordinate transformation;

[0009] Step 4: Reverse load the installation angle differences at different cross-sections of the primary hot state blade and the theoretical blade onto the theoretical airfoil, and at the same time adjust the geometric inlet and outlet angles to obtain the primary cold state blade;

[0010] Step 5: Conduct a pneumatic experiment on the primary cold state blade to obtain pneumatic results. Load the theoretical blade surface load, the pneumatic results of the primary cold state blade, and the centrifugal load onto the primary cold state blade through a CFD solver to obtain the secondary hot state blade, and deduce the installation angle, geometric inlet angle, and outlet angle of the secondary hot state blade from the deformation amount;

[0011] Step 6: Compare the installation angle, geometric inlet angle, and outlet angle of the theoretical blade and the secondary hot state blade, and judge whether the convergence criterion is met. When the convergence criterion is met, judge whether the performance requirements are satisfied. When the performance requirements are satisfied, obtain the processing data of the blade.

[0012] Preferably, the process of static analysis includes: designing a theoretical blade, and using a CFD solver to obtain the pneumatic load of the theoretical blade under set boundary conditions, and obtaining the centrifugal load based on the centrifugal force.

[0013] Preferably, the theoretical blade is the first-stage rotor blade of an axial compressor.

[0014] Preferably, the design conditions are the design point conditions at 100%, 80%, and 60% of the compressor design speed, corresponding to supersonic, transonic, and subsonic flow fields in the flow field respectively.

[0015] Preferably, the method for achieving the convergence criterion includes: comparing the differences in the installation angle, geometric inlet angle, and outlet angle at each cross-section between the secondary hot state blade and the theoretical blade. If the installation angle difference, geometric inlet angle difference, and outlet angle difference are within the specified error and meet the internal pneumatic performance of the compressor, it is considered that the convergence criterion is met; if the convergence criterion is not met, return to Step 3, and reverse load the characteristic angle differences at each cross-section between the secondary hot state blade and the theoretical blade onto the primary cold state blade to analyze the blade deformation amount again until the characteristic angles reach the convergence criterion.

[0016] Preferably, the method for meeting the performance requirements includes: modifying the fluid domain grid of the hot blade, performing three-dimensional CFD calculation and analysis at the designed rotational speed to obtain the aerodynamic performance and flow field characteristics of the compressor at the designed rotational speed, determining whether the performance requirements are met. If the requirements are met, the blade profile data of the secondary hot blade is used as the final blade profile data; if not, the error is further reduced, and the process returns to step three to iterate according to the new convergence conditions. Through repeated iteration of the above steps, the processing data of the compressor blade that finally meets the performance requirements is obtained.

[0017] The present invention discloses the following technical effects:

[0018] A blade anti-twist design method based on blade section displacement - characteristic angle proposed by the present invention comprehensively considers the influence of aerodynamic load and centrifugal load on blade deformation during the operation of the gas turbine compressor blade, effectively solves the blade anti-twist design problem of the compressor at subsonic, transonic and supersonic speeds, and improves the aerodynamic performance of the compressor. And the spatial displacement variable of the hot blade profile relative to the theoretical design blade profile at different sections along the blade height direction is converted, and the angle value transformation is carried out in the parameter form of geometric characteristic angles (including the blade profile installation angle, geometric inlet and outlet angles of the section). On the basis of ensuring the surface of the compressor blade, the surrounding flow conditions and the overall aerodynamic performance, the blade anti-twist design efficiency can be greatly improved, and the design cycle of the gas turbine can be shortened. BRIEF DESCRIPTION OF THE DRAWINGS

[0019] In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required for use in the embodiments. Obviously, the drawings described below are only some embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other drawings can be obtained based on these drawings.

[0020] Figure 1 It is a schematic flow chart of the method of the embodiment of the present invention. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0021] The following will clearly and completely describe the technical solutions in the embodiments of the present invention with reference to the drawings in the embodiments of the present invention. Obviously, the described embodiments are only some embodiments of the present invention, rather than all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts belong to the scope of protection of the present invention.

[0022] To make the above objects, features and advantages of the present invention more obvious and understandable, the present invention will be further described in detail below with reference to the drawings and specific embodiments.

[0023] AsFigure 1 As shown in the figure, the present invention provides a blade anti-twist design method based on blade cross-section displacement - characteristic angle, and the specific implementation manner of the method is achieved through the following steps:

[0024] Step 1: Conduct a static analysis under the action of aerodynamic load and centrifugal load in the design condition. Firstly, during iteration, assume the theoretical designed blade is AS, the aerodynamic load is Ps, the installation angle is θs, and the geometric inlet angle and outlet angle are β 1s and β 2s . Since the centrifugal load of the blade is determined by the centrifugal force generated by rotation, the centrifugal load is fixed during the entire iteration process. Use the CFD solver to obtain the aerodynamic load of the theoretical blade under the set boundary conditions.

[0025] Step 2: Considering the non-linear variable stiffness characteristics of the blade, use the structural mechanics solver to analyze and calculate the external force vector under the combined action of the aerodynamic load and centrifugal load obtained by the theoretical blade under the set boundary conditions, and combine with the internal force vector of the theoretical designed blade to obtain the unbalanced load. When the internal and external force loads of the blade reach equilibrium, the primary hot-state deformation amount IS can be obtained.

[0026] Step 3: Extract the deformation amounts of multiple cross-sections of the theoretical blade profile along the blade height direction, including the top and bottom cross-sections of the blade, and obtain the primary hot-state blade profile through coordinate transformation. Analyze the installation angle θ hot1 , geometric inlet angle β 1hot1 and outlet angle β 2hot1 for all blade profiles, and compare the differences between them.

[0027] Step 4: Reverse load the installation angle difference at different cross-sections between the primary hot-state blade and the theoretical designed blade onto the theoretical designed blade profile, and at the same time adjust the geometric inlet and outlet angles to form the primary cold-state blade profile. At the same time, each blade profile corresponds to different installation angles θ cold1 , geometric inlet angle β 1cold and outlet angle β 2cold .

[0028] Furthermore, optimize the scheme, obtain the spatial position information of the grid nodes at each cross-section of the working blade to obtain the preliminary state of the blade, load the aerodynamic load and centrifugal load onto the theoretical working blade respectively to obtain the spatial displacement vector set at each cross-section, and fit the change amounts of the installation angle, geometric inlet angle and outlet angle at each cross-section based on the new spatial position deviation vector set for anti-twist to obtain the first cold-state blade.

[0029] Step 5: Since the error is relatively large when loading the theoretical designed blade surface load under the corresponding boundary conditions of the CFD solver onto the calculated deformation of the primary cold-state blade, the aerodynamic results of the primary cold-state blade under the same boundary conditions are loaded during the calculation, and the centrifugal load is loaded simultaneously to obtain the secondary hot-state blade. At the same time, the installation angle θ is deduced from the deformation amount at the corresponding cross-section. hot2 , the geometric inlet angle β 1hot2 and the outlet angle β 2hot2 .

[0030] Step 6: Compare the differences in the installation angles and the inlet and outlet geometric angles at each cross-section between the secondary hot-state blade and the theoretically designed blade. If the differences in the installation angles and the inlet and outlet geometric angles are within the error range and meet the internal aerodynamic performance of the compressor, it is considered that convergence is achieved, and this analysis process can be terminated; otherwise, return to Step 3, and reverse load the characteristic angle differences at each cross-section between the secondary hot-state blade and the theoretical blade onto the primary cold-state blade to analyze the blade deformation amount again. Repeat this iterative process until the characteristic angles meet the convergence criterion.

[0031] Furthermore, for the optimized solution, the convergence condition for the reverse twist design: if the difference in the installation angle is within 0.1° and the difference in the inlet and outlet geometric angles is within 0.2°.

[0032] Step 7: Modify the fluid domain mesh of the hot-state blade profile, perform three-dimensional CFD calculation and analysis at the design speed, obtain the aerodynamic performance and flow field characteristics of the compressor at the design speed, and judge whether the performance requirements are met. If the requirements are met, use the blade profile data obtained in Step 6 as the final blade profile data; if not, further narrow the error condition in Step 6 and return to Step 6 to perform iteration according to the new convergence condition. Through the repeated iteration of the above steps, the processing data of the compressor blade that finally meets the performance requirements is obtained.

[0033] In summary, the significant advantages of the present invention compared with the prior art are summarized as follows:

[0034] 1) A blade reverse twist design method based on blade cross-section displacement-characteristic angle proposed by the present invention comprehensively considers the influence of the aerodynamic load and centrifugal load acting on the blade deformation during the operation of the gas turbine compressor blade, effectively solves the blade reverse twist design problem of the compressor under subsonic, transonic and supersonic conditions, and improves the aerodynamic performance of the compressor.

[0035] 2) This method converts the spatial displacement variables of the hot blade profiles at different cross-sections along the blade height direction relative to the theoretical designed blade profiles, and conducts angular value transformation in the form of parameters of geometric feature angles (including the installation angle of the cross-section blade profile, geometric inlet and outlet angles). On the basis of ensuring the blade surface, the surrounding flow conditions and the overall aerodynamic performance of the compressor, the blade reverse-twist design efficiency can be greatly improved, and the design cycle of the gas turbine can be shortened.

[0036] 3) This method can also be an important method for the practical engineering application of blade reverse-twist design in the fields of aviation and shipbuilding.

[0037] The embodiments described above are only used to describe the preferred mode of the present invention, rather than to limit the scope of the present invention. Without departing from the design spirit of the present invention, various deformations and improvements made by those of ordinary skill in the art to the technical solution of the present invention shall fall within the protection scope determined by the claims of the present invention.

Claims

1. A blade anti-twist design method based on blade cross-section displacement - characteristic angle, characterized in that, it includes the following steps: Step 1: Conduct a static analysis of the theoretical blade under the action of aerodynamic load and centrifugal load at the design condition; Step 2: Calculate the external force vector through a structural mechanics solver based on the aerodynamic load and centrifugal load, and obtain the primary hot-state deformation amount when the internal force vector and external force vector of the theoretical blade are balanced; Step 3: Extract the hot-state deformation amounts of multiple cross-sections of the theoretical blade along the blade height direction, and obtain the primary hot-state blade through coordinate transformation; Step 4: Reverse-load the installation angle differences at different cross-sections of the primary hot-state blade and the theoretical blade onto the theoretical airfoil, and at the same time adjust the geometric inlet and outlet angles to obtain the primary cold-state blade; Step 5: Conduct an aerodynamic experiment on the primary cold-state blade to obtain aerodynamic results. Load the surface load of the theoretical blade, the aerodynamic results of the primary cold-state blade, and the centrifugal load onto the primary cold-state blade through a CFD solver to obtain the secondary hot-state blade, and deduce the installation angle, geometric inlet angle, and outlet angle of the secondary hot-state blade from the deformation amount; Step 6: Compare the installation angle, geometric inlet angle, and outlet angle of the theoretical blade and the secondary hot-state blade, judge whether the convergence criterion is met. When the convergence criterion is met, judge whether the performance requirements are satisfied. When the performance requirements are satisfied, obtain the processing data of the blade.

2. The blade anti-twist design method based on blade cross-section displacement - characteristic angle according to claim 1, characterized in that, the process of the static analysis includes: designing the theoretical blade, and using a CFD solver to obtain the aerodynamic load of the theoretical blade under the set boundary conditions, and obtaining the centrifugal load based on the centrifugal force.

3. The blade anti-twist design method based on blade cross-section displacement - characteristic angle according to claim 2, characterized in that, the theoretical blade is the first-stage rotor blade of an axial-flow compressor.

4. The blade anti-twist design method based on blade cross-section displacement - characteristic angle according to claim 1, characterized in that, the design condition is the design point condition at 100%, 80%, and 60% of the compressor design speed, corresponding to supersonic, transonic, and subsonic flow field flows in the flow field respectively.

5. The blade anti-twist design method based on blade cross-section displacement - characteristic angle according to claim 1, characterized in that, the method for reaching the convergence criterion includes: comparing the differences in the installation angle, geometric inlet angle, and outlet angle at each cross-section of the secondary hot-state blade and the theoretical blade. If the installation angle difference, geometric inlet angle difference, and outlet angle difference are within the specified error and meet the internal aerodynamic performance of the compressor, it is considered that the convergence criterion is reached; if the convergence criterion is not reached, return to Step 3, and reverse-load the characteristic angle differences at each cross-section of the secondary hot-state blade and the theoretical blade onto the primary cold-state blade to analyze the blade deformation amount again until the characteristic angle reaches the convergence criterion.

6. The blade anti-twist design method based on blade cross-section displacement - characteristic angle according to claim 5, characterized in that, The method for meeting the performance requirements includes: modifying the fluid domain grid of the hot blade, performing three-dimensional CFD calculation and analysis at the design speed to obtain the aerodynamic performance and flow field characteristics of the compressor at the design speed, judging whether the performance requirements are met. If the requirements are met, the airfoil data of the secondary hot blade is used as the final airfoil data; if not, the error is further reduced, and the process returns to step three to perform iteration according to the new convergence conditions. Through repeated iteration of the above steps, the machining data of the compressor blade that finally meets the performance requirements is obtained.

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