A large-scale adaptive blade parameterized finite element direct modeling method

Abstract

The application relates to the field of blade structure parameterization design and modeling analysis, and provides a large self-adaptive blade parameterization finite element direct modeling method, which comprises the following steps: first, establishing section airfoil data points according to blade shape parameters and airfoil data; second, determining an occupying airfoil section in combination with structure layout parameters and shape parameters; third, introducing bias to generate a self-adaptive blade space occupying section set according to blade shape parameters; fourth, generating blade grid nodes in the longitudinal direction according to corresponding control points of each occupying airfoil; fifth, generating a skin and web element set; and sixth, carrying out block division and attribute assignment of the grid elements according to blade structure forms and layer parameters, forming a finite element analysis model, and exporting a formatted finite element calculation file. The application directly generates a high-quality blade finite element model by using airfoil parameters, blade design parameters, structure / attribute / grid parameters and the like.

Description

Technical Field

[0001] This application belongs to the field of parametric design and modeling analysis of blade structures, and specifically relates to a rapid and high-quality modeling method for adaptive large composite blades with pre-bending / sweep. Background Technology

[0002] With the increasing demand for renewable energy, the length of large wind turbine blades has exceeded 100 meters. Due to the rapid increase in blade load, the design of large adaptive (pre-bent / swept) blades has become an important optimization direction. On the one hand, the modification of the geometric shape in adaptive blade design leads to repeated modeling during the analysis process, resulting in low design iteration efficiency; on the other hand, the layup design of large composite blades is complex, and completing the structural finite element modeling requires processing a large number of material property parameters, resulting in a huge workload for analysis iteration. This application addresses the problems existing in the design and analysis of the aforementioned large blades by proposing to use parametric modeling to automatically generate the model. Starting from the airfoil data points, a finite element model is directly generated without the need for blade geometry and structural modeling. Summary of the Invention

[0003] The purpose of this application is to propose a direct parametric finite element modeling method for large adaptive blades. By utilizing airfoil data, combined with blade shape, structure, material properties, and mesh generation parameters, a high-quality finite element model can be quickly established, improving the efficiency of blade design, analysis, and optimization. Furthermore, it can provide a rapid modeling tool for the design of adaptive blades.

[0004] The technical solution adopted in this invention is as follows:

[0005] A direct modeling method for parametric finite element method of large adaptive blades includes:

[0006] A direct modeling method for parametric finite element method of large adaptive blades, characterized by the following steps:

[0007] Step 1: Determine the spaced section airfoil data points based on the original airfoil data points;

[0008] Step 2: Generate control points for the spaced airfoil section;

[0009] Step 3: Generate an adaptive blade space occupancy section set using offsetting.

[0010] Step 4: Generate the blade skin node set and finite element elements;

[0011] Step 5: Generate blade web node set and finite element elements;

[0012] Step 6: Assign parametric attributes and output the finite element model file.

[0013] Furthermore, in step two, based on the blade structure layout parameters and combined with the blade shape parameters, discrete interpolation / positioning is used to generate the main occupying airfoil control points; further, segmented occupying airfoil control points are generated sequentially in segments; thereby forming the occupying airfoil section control points.

[0014] Furthermore, in step three, the offset of each of the space-occupying airfoil section control points is determined based on the blade axis pre-bending or sweep parameters of the blade shape parameters, thus forming the adaptive blade space-occupying section set.

[0015] Furthermore, the sequence of generating segmented control points for the airfoil is as follows: segmented control points are generated sequentially for the upper trailing edge, upper rear skin, upper sparsity cap, upper front skin, leading edge, lower front skin, lower sparsity cap, lower rear skin, and lower trailing edge of the airfoil.

[0016] Furthermore, in step four, the meshing parameters are determined by fitting the spatial curves of the control points corresponding to each space occupancy section in the adaptive blade space occupancy section set, interpolating to generate longitudinal mesh nodes of the blade, forming a blade skin node set, and generating skin finite element elements along the longitudinal direction of the blade column by column.

[0017] Furthermore, in step five, based on the corresponding control points on the upper and lower sides of the blade skin, a web node set is generated by interpolation, and web finite element elements are generated preferentially according to the longitudinal columns.

[0018] Furthermore, in step six, the method for assigning attributes is as follows: based on the attribute parameters determined by the blade structure design, the finite element elements of the blade are divided into attribute blocks along the circumferential and longitudinal directions to form a circumferential-longitudinal 2D block set; based on the blade structure layout parameters and composite material layup parameters, parameterized layup settings are performed on each attribute block set of the blade, and a local coordinate system is established with the blade axis as the main axis to assign materials, layup and attributes.

[0019] Furthermore, in step one, the original airfoil data points are fitted and the interpolation points are densified using three partitioned spline curves to form the spaced-section airfoil data points.

[0020] Furthermore, in step two, the interpolation / positioning uses linear interpolation of discrete points in 2D space.

[0021] Furthermore, in steps five and six, the skin finite element element and the web finite element element are four-node rectangular shell elements.

[0022] The beneficial effects of this invention are:

[0023] 1. The method proposed in this invention starts from airfoil data points and directly generates a finite element model by skipping blade geometry and structural modeling. This method of directly generating a high-quality finite element analysis model based on low-dimensional airfoil data greatly simplifies the modeling steps and improves modeling efficiency.

[0024] 2. The method proposed in this invention is a direct finite element modeling method that relies on parameter input. It automatically generates a finite element mesh model through high-precision shape fitting and control point interpolation, resulting in a highly accurate model. The parameterized model makes the modification operation of the model more convenient and highly operable, and can accelerate the iteration speed when used for blade structure optimization.

[0025] 3. The proposed solution adopts parametric layup settings, which can quickly complete the definition of composite material layup that was originally complex and time-consuming, and further improve the efficiency of design and optimization of large wind turbine blades. Attached Figure Description

[0026] To more clearly illustrate the specific embodiments and technical solutions of the present invention, the accompanying drawings used in the existing implementation schemes or technical descriptions will be briefly introduced below.

[0027] Figure 1 This is a technical roadmap for a direct modeling method of parametric finite element method for large adaptive blades proposed in this application.

[0028] Figure 2 This is a schematic diagram of airfoil point spline fitting / interpolation and control node generation for a direct modeling method of parametric finite element method for large adaptive blades proposed in this application.

[0029] Figure 3 This is a schematic diagram of the partitioned airfoil cross-section structure of a large adaptive blade parametric finite element direct modeling method according to this application.

[0030] Figure 4 This is a schematic diagram of longitudinal node fitting / interpolation generation for a pre-bent / swept blade in a direct modeling method of parametric finite element method for large adaptive blades according to this application.

[0031] Figure 5 This is a schematic diagram of the generation of skin elements and web elements sequentially from blade mesh nodes in a direct modeling method of parametric finite element modeling for large adaptive blades according to this application.

[0032] Figure 6 The finite element model is established using the NREL5MW blade as input, with a pre-bending at 75% of the blade and a tip pre-bending angle of 5°, according to the parametric finite element direct modeling method of this application for large adaptive blades.

[0033] Figure 7The finite element model is established using the NREL5MW blade as input, with a 50% sweepback of the blade and a 10° tip sweepback angle, according to the parametric finite element direct modeling method of this application for large adaptive blades.

[0034] Figure 8 The method described in this application is a direct modeling method for parametric finite element analysis of large adaptive blades. Using the NREL5MW blade as input, the method generates attribute blocks (blade root) of the finite element model.

[0035] In the figure: 1. Original airfoil data points; 2. Airfoil data points of the occupancy section; 3. Main occupancy airfoil control points; 4. Segmented spline curve; 5. Main beam position; 6. Upper trailing edge of the airfoil; 7. Upper rear skin; 8. Upper beam cap; 9. Upper front skin; 10. Leading edge; 11. Lower front skin; 12. Lower beam cap; 13. Lower rear skin; 14. Lower trailing edge; 15. Segmented occupancy airfoil control points; 16. Adaptive blade spatial occupancy section set; 17. Spatial curve; 18. Skin mesh node set; 19. Y-direction offset; 20. X-direction offset; 21. Skin longitudinal; 22. Web finite element element; 23. Web vertical element; 24. Web longitudinal element; 25. Rectangular four-node element; 26. Skin chordal element; 27. Skin longitudinal element. Detailed Implementation Plan

[0036] To make the objectives, technical solutions, and advantages of this application clearer, the accompanying drawings used in the embodiments will be briefly described below. Obviously, the drawings described below are merely some embodiments of the present invention, and not all embodiments. The embodiments described below with reference to the accompanying drawings are exemplary and intended to explain this application, and should not be construed as limiting this application. Based on the embodiments in this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application. Examples of this application will be described in detail below with reference to the accompanying drawings.

[0037] In the description of this application, it should be understood that the terms "center," "longitudinal," "lateral," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," and "outer," etc., which indicate orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings, are only for the convenience of describing this application and simplifying the description, and do not indicate that the device or component referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limiting the scope of protection of this application.

[0038] The following is in conjunction with the appendix Figure 1 To be continued Figure 8 The specific implementation plan of this application is explained.

[0039] This embodiment provides a direct parametric finite element modeling method for large adaptive blades, which directly generates a parameter-driven finite element analysis model of the blade using airfoil data and blade parameters, such as... Figure 1 The diagram illustrates one implementation method of this application, and the main implementation process includes:

[0040] S1 fitting / interpolation determines high-quality spaced airfoil data points;

[0041] S2 generates the control points for the spaced airfoil section;

[0042] S3 bias generates an adaptive blade space occupancy section set;

[0043] S4 generates the blade skin node set and finite element elements;

[0044] S5 generates the blade web node set and finite element elements;

[0045] S6 attribute partitioning forms a partition set;

[0046] S7 assigns composite material layup properties to each block;

[0047] S8 outputs model data files.

[0048] The following is a detailed explanation of the specific operation of the above method, including the following detailed implementation steps:

[0049] S1, such as Figure 2 As shown, based on the original airfoil data point 1 and the blade chord length distribution parameters and twist angle distribution parameters of the blade shape parameters, high-quality spaced section airfoil data point 2 is determined by fitting / interpolation.

[0050] In the plane, the rough original airfoil data point 1 is fitted and the interpolation points are densified by three piecewise spline curves 4 (airfoil leading edge, airfoil upper rear part and airfoil lower rear part) to form the spaced section airfoil data point 2.

[0051] S2, such as Figure 3 As shown, based on the main beam position 5 (including the leading and trailing edge positions, etc.) of the blade structure layout parameters, combined with the blade chord length distribution parameters and twist angle distribution parameters of the blade shape parameters, discrete interpolation / positioning is used to generate the main occupying airfoil control points 3; further, segmented occupying airfoil control points 15 are generated sequentially; thus, complete occupying airfoil section control points are obtained to form the occupying airfoil section.

[0052] Specifically, in the plane, based on the encrypted spaced airfoil data points 2, the discrete interpolation method is used to accurately locate the main control points, and segmented control points are generated sequentially for the upper trailing edge 6, upper rear skin 7, upper beam cap 8, upper front skin 9, leading edge 10, lower front skin 11, lower beam cap 12, lower rear skin 13, and lower trailing edge 14 of the airfoil to form the spaced airfoil section.

[0053] S3, such as Figure 4 As shown, the space-occupying airfoil section is offset according to the blade axis pre-bending or sweep parameters to form an adaptive blade space-occupying section set 16.

[0054] Specifically, based on the curve type of the adaptive blade's pre-bending / sweep, such as a parabola or a power function curve, the X-direction offset 20 and Y-direction offset 19 (with the blade axis as the Z-axis) of each airfoil section control point are determined, forming an adaptive blade spatial occupancy section set 16. In one embodiment, the pre-bending / sweep adopts a single-parameter parabola, and the parameter can be the offset at the blade tip or the tangent angle.

[0055] S4, such as Figure 4 As shown, spatial curves are fitted based on the control points corresponding to each space-occupying section in the adaptive blade space-occupying section set 17; meshing parameters such as feature dimensions are determined according to the needs of model analysis, and combined with the above meshing parameters, longitudinal (Z direction) mesh nodes of the blade are generated by interpolation to form a blade skin mesh node set 18.

[0056] Furthermore, control points corresponding to each occupied section are selected, and spline fitting is performed in the longitudinal-vertical (ZY) and longitudinal-chordal (ZX) directions respectively to generate longitudinal spatial curves 17. Interpolation of the longitudinal grid nodes is then performed to form the grid node set of the blade skin. Custom node objects include attributes such as node.id, node spatial coordinates.coords, and the element set to which the node belongs. The skin node objects are distributed in two dimensions along the longitudinal and circumferential directions, with longitudinal nodes taking priority.

[0057] Furthermore, such as Figure 5 As shown, based on the skin mesh node set 18, the longitudinal skin element 27 is generated along the blade longitudinal direction (Z direction) in column priority 21, and the chordal skin element 26 is used to perform region division and block division.

[0058] Furthermore, the skinning unit adopts a custom object four-node matrix unit 25, including unit.id, unit sequential node set.nodes, attribute set of the unit.pset, and unit attribute definition coordinate system label.system, etc. The skinning unit objects are distributed in two dimensions along the vertical and circumferential directions, and are all four-node rectangular shell units.

[0059] S5, such as Figure 5As shown, based on the skin mesh node sets 18 corresponding to the upper and lower skin as control points, the web node set is generated by interpolation. Further, web finite element elements 22 are generated prioritizing vertical columns. The web elements use custom object four-node matrix elements 25. The method for generating node sets and finite element data is similar to the method for generating skin node sets and finite element data in S4.

[0060] S6. Based on the attribute parameters determined by the blade structure design, the finite element elements of the blade are divided into attribute blocks along the circumferential and longitudinal directions to form a two-dimensional block set of web vertical element 23 and web longitudinal element 24.

[0061] The skin is divided circumferentially into regions such as upper / lower rear edge reinforcement, upper / lower rear skin, upper / lower skin beam cap, upper / lower front skin, and skin front edge reinforcement. Along the longitudinal direction, regions are divided according to the skin composite material layup parameters, forming a circumferential-longitudinal 2D block set. Each block contains units with the same material properties; their unit numbers are recorded and defined as a property set.

[0062] S7. Based on the blade structure layout parameters and composite material layup parameters, parameterized layup settings are performed on the blade attribute block set. A local coordinate system is established with the blade axis as the main axis to assign materials, layup and attributes.

[0063] Furthermore, for adaptive blades, due to pre-bending / sweeping, a local coordinate system is determined by the tangent direction of the blade axis. The reference coordinate system for the composite material layup is the corresponding local coordinate system. A composite material layup with a fixed layup ratio is used for layup design, and the layup thickness is assigned and controlled by parameters in the longitudinal and circumferential directions.

[0064] S8. Output the formatted calculation file according to the set solver to form the finite element model of the blade structure.

[0065] The calculation file output by the finite element solver contains corresponding keyword data blocks, including node data, element data, material data, grouping data, local coordinate system data, and layup attribute data.

[0066] Figure 6 and Figure 7 These are different views of a pre-bent and swept blade model established using the method proposed in this invention. Figure 8 The method proposed in this invention demonstrates the grouping settings for composite material properties, with different colors representing different material property groups. It is evident that the finite element model of the blade directly established using this method has high accuracy and can generate effective finite element models of blades with pre-bending and sweepback. Furthermore, the method is simple to operate and can automate parameter settings for composite materials.

[0067] The above description of the disclosed embodiments enables those skilled in the art to make or use the invention. Various modifications to these embodiments will be readily apparent to those skilled in the art, and the general principles defined herein may be implemented in other embodiments without departing from the spirit or scope of the invention. Therefore, the invention is not to be limited to the embodiments shown herein, but is to be accorded the widest scope consistent with the principles and novel features disclosed herein.

Claims

1. A direct parametric finite element modeling method for large adaptive blades, characterized in that, Includes the following steps: Step 1: Determine the spaced section airfoil data points (2) based on the original airfoil data points (1); Step 2: Generate control points for the spaced airfoil section; Step 3: Based on the blade axis pre-bending or sweep parameters of the blade shape parameters, determine the X-direction offset and Y-direction offset of each of the aforementioned space-occupying airfoil section control points, wherein the blade axis is the Z-axis, and generate an adaptive blade space-occupying section set through the offsets; Step 4: Fit the spatial curves of the control points corresponding to each space occupancy section in the adaptive blade space occupancy section set, determine the mesh generation parameters, interpolate to generate longitudinal mesh nodes of the blade, form a blade skin node set, and generate skin finite element elements along the longitudinal direction of the blade in column priority. Step 5: Based on the corresponding control points on the upper and lower sides of the blade skin, interpolate to generate a web node set, and further generate web finite element elements in the order of longitudinal columns. Step 6: Assign parameterized attributes and output the finite element model file; The method does not establish a blade geometry and structure model, but directly generates a finite element model from the airfoil data points and the space-occupying section set; the finite element model file is a formatted calculation file oriented towards the finite element solver.

2. The method for direct parametric finite element modeling of large adaptive blades according to claim 1, characterized in that, In step two, based on the blade structure layout parameters and combined with the blade shape parameters, discrete interpolation / positioning is used to generate the main occupying airfoil control points (3), and further segmented segmented occupying airfoil control points (15) are generated sequentially to form the occupying airfoil section control points.

3. The method for direct parametric finite element modeling of large adaptive blades according to claim 2, characterized in that, The sequence of generating segmented control points for the airfoil is as follows: segmented control points are generated sequentially for the upper trailing edge (6), upper rear skin (7), upper beam cap (8), upper front skin (9), leading edge (10), lower front skin (11), lower beam cap (12), lower rear skin (13), and lower trailing edge (14) of the airfoil.

4. The method for direct parametric finite element modeling of large adaptive blades according to claim 3, characterized in that, In step six, the method for assigning attributes is as follows: based on the attribute parameters determined by the blade structure design, the finite element elements of the blade are divided into attribute blocks along the circumferential and longitudinal directions to form a two-dimensional block set in the circumferential and longitudinal directions; based on the blade structure layout parameters and composite material layup parameters, parameterized layup settings are performed on each attribute block set of the blade, and a local coordinate system is established with the blade axis as the main axis to assign materials, layup and attributes.

5. The method for direct parametric finite element modeling of large adaptive blades according to claim 4, characterized in that, In step one, the original airfoil data points (1) are fitted and the interpolation points are densified by three partitioned spline curves to form the spaced section airfoil data points (2).

6. The method for direct parametric finite element modeling of large adaptive blades according to claim 5, characterized in that, In step two, the interpolation / positioning uses linear interpolation of discrete points in two-dimensional space.

7. The method for direct parametric finite element modeling of large adaptive blades according to claim 6, characterized in that, In steps five and six, the skin finite element unit and the web finite element unit are four-node rectangular shell units.

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