A method for optimizing the layout of airfoil structures based on parametric modeling and finite element analysis

Abstract

This application belongs to the field of aircraft structural design and optimization technology, and specifically relates to a method for optimizing wing structure layout based on parametric modeling and finite element analysis. The method includes: Step 1, obtaining the aerodynamic geometric model of the wing structure layout scheme and preprocessing the aerodynamic geometric model; Step 2, parametrically partitioning the aerodynamic geometric model of the wing and extracting the geometric information of the partitioned elements; Step 3, performing finite element preprocessing based on the geometric information to obtain the wing finite element model; Step 4, performing finite element analysis based on the wing finite element model and obtaining the analysis results; Step 5, determining whether the analysis results meet the requirements, returning to Step 2 to re-perform parametric partitioning, and repeating Steps 2 to 5 for optimization iteration until the convergence condition is met, thereby obtaining the optimal wing structure layout scheme that meets the requirements.

Description

Technical Field

[0001] This application belongs to the field of aircraft structural design and optimization technology, and specifically relates to a method for optimizing the layout of wing structures based on parametric modeling and finite element analysis. Background Technology

[0002] Modern aircraft design is an extremely complex, multidisciplinary, and costly engineering process. As the core aerodynamic surface and main load-bearing structure of an aircraft, the design quality of the wing directly determines its performance, safety, and economy. In wing structural layout design, the internal layout—the number, spacing, and location distribution of ribs and longitudinal walls—is a key factor affecting structural weight, stiffness, strength, flutter characteristics, and maintainability.

[0003] Traditional wing structure layout design heavily relies on designers' experience and past solutions, typically employing a sequential, iterative, manual cycle of design-modeling-analysis-modification. This model has several inherent drawbacks: First, it is extremely inefficient. Every layout modification, whether adding a wing rib or moving the longitudinal wall, requires engineers to spend a significant amount of time on geometric modeling, cleanup, mesh generation, material property assignment, load and boundary condition settings, and other preprocessing work. This process involves a high proportion of repetitive labor, is lengthy, and severely restricts the speed of design iteration. Second, it severely limits the exploration of the design space. Constrained by time and manpower, engineers can usually only analyze and compare a limited number of candidate solutions, making it difficult to systematically explore the vast and non-intuitive design space. This means that the final solution may only be a locally optimal solution, rather than the globally lightest or best-performing solution, missing significant weight reduction opportunities. Furthermore, model consistency and accuracy are difficult to guarantee. In manual operation, especially when dealing with shared nodes of meshes at the junction of different design areas and when transferring loads and constraints between different components, human error is easily introduced, leading to distortion of the analysis model, unreliable calculation results, and consequently affecting the correctness of design decisions.

[0004] Although parametric modeling and optimization design theories have made significant progress in recent years, and some commercial software has emerged to assist in structural optimization, their applications are mostly concentrated on optimizing component dimensions such as skin thickness and beam flange area, or performing topology optimization based on a fixed internal layout. There is still a significant gap in existing technology for automating processes that directly optimize the layout of ribs and longitudinal walls—that is, the structure's topology and position. Most solutions fail to achieve seamless integration across the entire chain from the original aerodynamic shape to the final finite element model, and their capabilities in handling structural generation on complex surfaces, multi-region collaborative meshing, and automatically ensuring mesh quality are limited. In particular, there is a lack of an end-to-end closed-loop method that can tightly couple parametric geometric modeling, high-fidelity finite element analysis, automatic application of multi-condition loads, and intelligent optimization algorithms.

[0005] Therefore, there is an urgent need for a technical solution to overcome or mitigate at least one of the aforementioned defects in the existing technology. Summary of the Invention

[0006] The purpose of this application is to provide a method for optimizing the layout of airfoil structures based on parametric modeling and finite element analysis, so as to solve at least one problem existing in the prior art.

[0007] The technical solution of this application is:

[0008] A method for optimizing the layout of airfoil structures based on parametric modeling and finite element analysis includes:

[0009] Step 1: Obtain the aerodynamic geometric model of the wing structure layout scheme, and preprocess the aerodynamic geometric model of the wing.

[0010] Step 2: Parametrically mesh the aerodynamic geometric model of the wing and extract the geometric information of the meshed units;

[0011] Step 3: Perform finite element preprocessing based on the geometric information to obtain the finite element model of the wing;

[0012] Step 4: Perform finite element analysis based on the finite element model of the wing and obtain the analysis results;

[0013] Step 5: Determine whether the analysis results meet the requirements. Return to Step 2 to perform parametric subdivision again. Repeat Steps 2 to 5 for optimization iteration until the convergence condition is met and the optimal wing structure layout scheme that meets the requirements is obtained.

[0014] In at least one embodiment of this application, step one involves obtaining an aerodynamic geometry model of the wing's structural layout and preprocessing the aerodynamic geometry model, including:

[0015] Obtain the aerodynamic geometric model of the wing structure layout scheme, wherein the aerodynamic geometric model of the wing includes the wing skin surface;

[0016] Determine the root chord and tip chord of the aerodynamic geometry model of the wing, and construct a continuous chord plane;

[0017] The optimization range is determined on the chord plane, and the wing is divided into multiple independent optimization design regions along the spanwise and chordwise directions by the separation surface.

[0018] In at least one embodiment of this application, the wing is divided along its span into a root section optimization design region, a middle section optimization design region, and an outer section optimization design region.

[0019] In at least one embodiment of this application, step two involves parametrically subdividing the aerodynamic geometry model of the wing and extracting the geometric information of the subdivided units, including:

[0020] The number of ribs, the spanwise position of the ribs, the number of longitudinal walls, and the chordwise position of the longitudinal walls in each of the optimized design regions are defined as optimization variables;

[0021] In each of the optimized design regions, wing rib positioning lines and longitudinal wall positioning lines perpendicular to the chord plane are generated, and the wing skin surface is subdivided to obtain different subdivided units.

[0022] Extract the geometric information of the subdivided unit.

[0023] In at least one embodiment of this application, the segmentation unit includes a skin unit, a web unit, and a flange unit.

[0024] In at least one embodiment of this application, the geometric information includes coordinates, length, and area.

[0025] In at least one embodiment of this application, step three, which involves performing finite element preprocessing based on the geometric information, includes:

[0026] Mesh elements are generated based on the geometric information, wherein the skin mesh elements and the web mesh elements adopt quadrilateral shell meshes, and the flange mesh elements adopt beam element meshes.

[0027] Assign corresponding cell properties to different types of mesh cells;

[0028] Fixed supports are applied to all nodes at the wing root, transforming typical load conditions into nodal forces or pressures applied to the skin mesh elements.

[0029] In at least one embodiment of this application, the typical load conditions include the maximum overload load condition, the maximum negative overload load condition, the symmetrical maneuver load condition, and the rolling maneuver load condition.

[0030] In at least one embodiment of this application, in step four, the analysis results include at least the global maximum deformation, element stress level, and structural weight.

[0031] In at least one embodiment of this application, in step five, a wing structure layout scheme with the minimum structural weight that satisfies the stiffness and strength requirements under all typical load conditions is obtained.

[0032] The invention has at least the following beneficial technical effects:

[0033] The wing structure layout optimization method based on parametric modeling and finite element analysis proposed in this application can efficiently, automatically, and reliably achieve intelligent optimization of wing structure layout. This method is beneficial for improving aircraft design capabilities, shortening the R&D cycle, reducing manufacturing costs, and achieving significant weight reduction benefits. It has urgent industrial demand and significant practical significance. Attached Figure Description

[0034] Figure 1 This is a flowchart of an implementation method for wing structure layout optimization based on parametric modeling and finite element analysis according to one embodiment of this application;

[0035] Figure 2 This is a schematic diagram of the aerodynamic geometry of an airfoil according to one embodiment of this application;

[0036] Figure 3 This is a schematic diagram of the optimized design region division of one embodiment of this application;

[0037] Figure 4 This is a schematic diagram of finite element mesh generation according to one embodiment of this application;

[0038] Figure 5 This is a schematic diagram illustrating the application of constraints and loads in one embodiment of this application. Detailed Implementation

[0039] To make the objectives, technical solutions, and advantages of this application clearer, the technical solutions in the embodiments of this application will be described in more detail below with reference to the accompanying drawings. In the drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments are some, but not all, embodiments of this application. 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. All other embodiments obtained by those skilled in the art based on the embodiments of this application without creative effort are within the scope of protection of this application. The embodiments of this application will be described in detail below with reference to the accompanying drawings.

[0040] 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", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element 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.

[0041] The following is in conjunction with the appendix Figures 1 to 5This application will be described in further detail.

[0042] This application provides a method for optimizing the layout of airfoil structures based on parametric modeling and finite element analysis, including the following steps:

[0043] Step 1: Obtain the aerodynamic geometric model of the wing structure layout scheme and preprocess the aerodynamic geometric model of the wing.

[0044] Step 2: Parametrically mesh the aerodynamic geometric model of the wing and extract the geometric information of the meshed units;

[0045] Step 3: Perform finite element preprocessing based on geometric information to obtain the finite element model of the wing;

[0046] Step 4: Perform finite element analysis based on the finite element model of the wing and obtain the analysis results;

[0047] Step 5: Determine whether the analysis results meet the requirements. Return to Step 2 to perform parametric subdivision again. Repeat Steps 2 to 5 for optimization iteration until the convergence condition is met and the optimal wing structure layout scheme that meets the requirements is obtained.

[0048] The wing structure layout optimization method based on parametric modeling and finite element analysis proposed in this application includes, in step one, obtaining the wing aerodynamic geometry model of the wing structure layout scheme, and preprocessing the wing aerodynamic geometry model, including:

[0049] Obtain the aerodynamic geometry model of the wing structure layout scheme, which includes the wing skin surface;

[0050] Determine the root chord and tip chord of the wing's aerodynamic geometry model, and construct a continuous chord plane;

[0051] The optimization range is determined on the chord plane, and the wing is divided into multiple independent optimization design regions along the spanwise and chordwise directions by the separation surface.

[0052] In the preferred embodiment of this application, such as Figure 1 As shown, firstly, geometric preprocessing is performed. The specific process is as follows:

[0053] S1.1 Importing the Wing Aerodynamic Shape: A Python script is used to read the initial geometric model of the wing's aerodynamic shape, such as IGES or STEP format files. This model includes surfaces describing the upper and lower surfaces of the wing skin. The imported surfaces are fitted and repaired to ensure their geometric integrity, providing a foundation for subsequent operations. Figure 2 A geometric model of a delta wing with linear chord length variation is presented.

[0054] S1.2 Establishing a Chord Plane Reference Frame: Based on the imported wing aerodynamic shape, a continuous chord plane is constructed by automatically calculating the root chord and tip chord of the wing using a Python script or by having the user specify the chord. This chord plane is an abstract reference plane, and its normal direction will serve as the directional reference for all subsequently generated ribs and longitudinal wall webs, ensuring that internal structural elements are perpendicular to the aerodynamic chord lines and comply with engineering design specifications.

[0055] S1.3 Define the optimization design region: On the chord plane, the optimization range is interactively specified using a Python script, such as optimizing only the central wing box. Furthermore, by introducing a separation surface, the wing is divided into multiple independent optimization design regions along the spanwise and chordwise directions. Figure 3 As shown. For example, the wing can be divided along its span into a root section optimization design region, a middle section optimization design region, and an outer section optimization design region, allowing different layout density rules to be used in different regions.

[0056] The wing structure layout optimization method based on parametric modeling and finite element analysis in this application, in step two, involves parametrically subdividing the wing's aerodynamic geometric model and extracting the geometric information of the subdivided elements, including:

[0057] The number of ribs, spanwise position of ribs, number of longitudinal walls, and chordwise position of longitudinal walls in each optimized design area are defined as optimization variables;

[0058] In each optimization design region, rib and longitudinal wall positioning lines perpendicular to the chord plane are generated, and the wing skin surface is subdivided to obtain different subdivided elements.

[0059] Extract the geometric information of the subdivided elements.

[0060] In a preferred embodiment of this application, the parametric modeling process is as follows:

[0061] S2.1 Define optimization variables: Define the number of ribs, spanwise position of ribs, number of longitudinal walls, and chordwise position of longitudinal walls in each optimization design area as core optimization variables, and determine the initial values ​​of the optimization variables.

[0062] S2.2 Perform parametric meshing: Using a Python script, wing rib positioning lines and longitudinal wall positioning lines are automatically generated in each optimization design region based on the current values ​​of these optimization variables. The original wing skin surface is then meshed, resulting in skin elements, web elements, and flange elements.

[0063] S2.3 Extract element geometry information: Then extract the geometry information of all subdivided elements, such as coordinates, length, area, etc.

[0064] The wing structure layout optimization method based on parametric modeling and finite element analysis in this application includes, in step three, finite element preprocessing based on geometric information, including:

[0065] Mesh elements are generated based on geometric information, wherein the skin mesh elements and the web mesh elements adopt quadrilateral shell meshes, and the flange mesh elements adopt beam element meshes;

[0066] Assign corresponding cell properties to different types of mesh cells;

[0067] Fixed supports are applied to all nodes at the wing root, transforming typical load conditions into nodal forces or pressures applied to the skin mesh elements.

[0068] In a preferred embodiment of this application, the specific process of finite element preprocessing is as follows:

[0069] S3.1 Mesh Generation: Based on the extracted geometric information, high-quality mesh elements are automatically generated, such as... Figure 4 As shown. Mesh cell generation follows these rules:

[0070] a. The default grid size is set to 1 / 15 of the local chord length. Users can override this default value globally or in specific areas.

[0071] b. At the junctions of different optimization design areas, and at the connections between ribs, longitudinal walls and skin, strictly ensure the common nodes of the mesh to ensure the correct transmission of forces and the accuracy of calculation.

[0072] c. The skin mesh element and the web mesh element adopt a quadrilateral shell mesh, and the edge mesh element adopts a beam element mesh.

[0073] S3.2 Assigning Attributes: Automatically assign corresponding Nastran attribute cards to different types of grid cells using Python scripts.

[0074] For skin mesh elements and web mesh elements, assign PSHELL integral shell property cards and MAT1 material property cards, and support PCOMP composite material layup property cards.

[0075] For edge mesh elements, assign attribute cards to PBAR (equal cross-section beam), PBEAM (variable cross-section beam), or PROD (bar) based on the cross-sectional shape, and associate them with the corresponding MAT1 material attribute cards.

[0076] S3.2 Applying constraints and loads:

[0077] like Figure 5As shown, fixed constraints are applied to all nodes at the wing root using the SPC1 single-point constraint set, restricting all six degrees of freedom. The wing root is typically located at the first wing rib. Using an engineering simplification method based on aerodynamic pressure distribution and inertial loads, four typical load cases for the wing are calculated. These typical load cases are then converted into nodal forces or pressures and automatically applied to the corresponding skin mesh elements using Nastran's FORCE concentrated force load card and PLOAD4 distributed force load card. The typical load cases include the maximum overload load case, the maximum negative overload load case, the symmetric maneuver load case, and the roll maneuver load case.

[0078] The wing structure layout optimization method based on parametric modeling and finite element analysis in this application, in step four, the specific process of finite element solution and post-processing is as follows:

[0079] A complete Nastran input file is automatically generated using a Python script, and the Nastran solver is called to perform static analysis. After the analysis is complete, the Nastran output file is parsed to extract key analysis results, including but not limited to global maximum deformation, element stress levels, and structural weight. It is also determined whether the element stress levels exceed allowable stresses.

[0080] The wing structure layout optimization method based on parametric modeling and finite element analysis proposed in this application, in step five, involves optimization iteration and scheme decision-making. Specifically, the obtained stiffness, strength, and weight results are fed back to the optimization algorithm, such as gradient descent or genetic algorithm. The optimization variables in step two are adjusted to meet the stiffness and strength requirements under all typical load conditions as constraints, with minimizing structural weight as the objective function. Steps two through five are executed iteratively through the optimization algorithm until the wing structure layout scheme with the minimum structural weight that meets the stiffness and strength requirements under all typical load conditions is found, and this scheme is taken as the final design output.

[0081] This application presents a wing structure layout optimization method based on parametric modeling and finite element analysis. By constructing an automated closed-loop system driven globally by Python scripts, with parametric modeling as the core and Nastran as the high-precision analysis engine, it achieves intelligent design throughout the entire process, from importing the wing aerodynamic shape to generating the final optimal wing structure layout. Its core objective is to automate the entire process of geometric modeling, mesh generation, attribute assignment, load application, and finite element analysis. Using the number and location of ribs and longitudinal walls as direct optimization variables, and with the goal of minimizing structural weight while constrained by stiffness and strength requirements, the method automatically iterates to find the lightest wing structure layout that satisfies all performance indicators efficiently and reliably. This significantly improves design quality and efficiency, achieving weight reduction benefits that are difficult to achieve with traditional methods. This application enables highly automated and integrated wing structure layout optimization, overcoming the shortcomings of existing technologies that rely on manual experience, have low design efficiency, and struggle to obtain globally optimal solutions.

[0082] The wing structure layout optimization method based on parametric modeling and finite element analysis proposed in this application has the following beneficial effects:

[0083] High degree of automation: It realizes full-process automation from geometry to analysis, which greatly reduces manual intervention and repetitive work and improves design efficiency;

[0084] Global optimization: By systematically exploring the design space through optimization algorithms, a better solution than traditional empirical design can be found to achieve the weight reduction goal;

[0085] High reliability: The programmed processing ensures the consistency of the model, especially the common nodes of the mesh, eliminates human error, and improves the credibility of the finite element analysis results;

[0086] High flexibility: Parametric design makes it very convenient to modify the scheme and easily adapt to different design requirements and the development of new models.

[0087] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. A method for optimizing the layout of an airfoil structure based on parametric modeling and finite element analysis, characterized in that, include: Step 1: Obtain the aerodynamic geometric model of the wing structure layout scheme, and preprocess the aerodynamic geometric model of the wing. Step 2: Parametrically mesh the aerodynamic geometric model of the wing and extract the geometric information of the meshed units; Step 3: Perform finite element preprocessing based on the geometric information to obtain the finite element model of the wing; Step 4: Perform finite element analysis based on the finite element model of the wing and obtain the analysis results; Step 5: Determine whether the analysis results meet the requirements. Return to Step 2 to perform parametric subdivision again. Repeat Steps 2 to 5 for optimization iteration until the convergence condition is met and the optimal wing structure layout scheme that meets the requirements is obtained.

2. The wing structure layout optimization method based on parametric modeling and finite element analysis according to claim 1, characterized in that, In step one, the aerodynamic geometric model of the wing structure layout scheme is obtained, and the aerodynamic geometric model of the wing is preprocessed, including: Obtain the aerodynamic geometric model of the wing structure layout scheme, wherein the aerodynamic geometric model of the wing includes the wing skin surface; Determine the root chord and tip chord of the aerodynamic geometry model of the wing, and construct a continuous chord plane; The optimization range is determined on the chord plane, and the wing is divided into multiple independent optimization design regions along the spanwise and chordwise directions by the separation surface.

3. The wing structure layout optimization method based on parametric modeling and finite element analysis according to claim 2, characterized in that, The wing is divided along its span into the root section optimization design region, the middle section optimization design region, and the outer section optimization design region.

4. The wing structure layout optimization method based on parametric modeling and finite element analysis according to claim 3, characterized in that, In step two, the aerodynamic geometric model of the wing is parametrically partitioned, and the geometric information of the partitioned units is extracted, including: The number of ribs, the spanwise position of the ribs, the number of longitudinal walls, and the chordwise position of the longitudinal walls in each of the optimized design regions are defined as optimization variables; In each of the optimized design regions, wing rib positioning lines and longitudinal wall positioning lines perpendicular to the chord plane are generated, and the wing skin surface is subdivided to obtain different subdivided units. Extract the geometric information of the subdivided unit.

5. The wing structure layout optimization method based on parametric modeling and finite element analysis according to claim 4, characterized in that, The segmented unit includes a skin unit, a web unit, and a flange unit.

6. The wing structure layout optimization method based on parametric modeling and finite element analysis according to claim 5, characterized in that, The geometric information includes coordinates, length, and area.

7. The wing structure layout optimization method based on parametric modeling and finite element analysis according to claim 6, characterized in that, Step three involves performing finite element preprocessing based on the geometric information, including: Mesh elements are generated based on the geometric information, wherein the skin mesh elements and the web mesh elements adopt quadrilateral shell meshes, and the flange mesh elements adopt beam element meshes. Assign corresponding cell properties to different types of mesh cells; Fixed supports are applied to all nodes at the wing root, transforming typical load conditions into nodal forces or pressures applied to the skin mesh elements.

8. The wing structure layout optimization method based on parametric modeling and finite element analysis according to claim 7, characterized in that, The typical load conditions include the maximum overload load condition, the maximum negative overload load condition, the symmetrical maneuver load condition, and the rolling maneuver load condition.

9. The wing structure layout optimization method based on parametric modeling and finite element analysis according to claim 8, characterized in that, In step four, the analysis results include at least the global maximum deformation, element stress level, and structural weight.

10. The wing structure layout optimization method based on parametric modeling and finite element analysis according to claim 9, characterized in that, In step five, the wing structure layout scheme with the minimum structural weight that meets the stiffness and strength requirements under all typical load conditions is obtained.

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