A method and apparatus for modeling a composite aircraft structure

Abstract

The application discloses a kind of composite aircraft structure modeling method and equipment, belong to aircraft structure computer aided engineering modeling technical field, to solve the problem of low precision, modification is complicated, the low efficiency of existing modeling method.The method comprises: constructing the two-dimensional geometric topology of aircraft component and shell element grid;According to the relative position relationship of structure, set the unit normal topology of component shell element;According to this topology, construct the lay-up sequence topology;Based on lay-up sequence topology, generate the standardized full-machine lay-up data table containing lay-up name, sequence, angle, thickness and unit set identification;Finally, the data table is input into CAE software, and the full-machine composite structure lay-up model is automatically generated.The application realizes high-precision expression and rapid global modification of model by combining topology definition with data table, significantly improves the efficiency and iteration speed of composite aircraft structure design analysis.

Description

Technical Field

[0001] This invention relates to the field of computer-aided engineering (CAE) modeling technology for aircraft structures, and in particular to a method and equipment for modeling composite material aircraft structures. It is especially applicable to the design and analysis stage of aircraft widely using all-composite material structures, such as skin, beams, frames, and ribs, in the low-altitude economic field, such as electric vertical takeoff and landing (eVTOL) aircraft and general aviation aircraft. Background Technology

[0002] With the rapid development of the low-altitude economy, the structural design of low-altitude eVTOL aircraft, general aviation aircraft, and other similar vehicles is increasingly trending towards the use of all-composite material structures to achieve lightweight, high strength, and high design freedom. In the design, analysis, and optimization iteration process of such aircraft, how to efficiently and accurately establish and manage the entire aircraft's composite material structure layup model in CAE software has become a key aspect of improving R&D efficiency.

[0003] Currently, there are two main methods in the industry for finite element modeling of aircraft composite material structures, but both have significant defects and shortcomings: 1. Equivalent Isotropic Material Modeling Method: This method treats composite materials with significant anisotropic properties as isotropic materials during modeling. This method fails to consider the anisotropic characteristics of composite materials and cannot account for coupling effects. Therefore, structural strength, stiffness, and dynamic analyses based on this simplified model have low accuracy, and the calculated results deviate significantly from actual physical behavior, failing to meet the requirements of high-reliability aircraft structural design.

[0004] 2. Region-based ply modeling: This method divides a complete composite component, such as an aircraft wing skin, into several distinct "regions" within its plane. Each region is assumed to have a completely consistent ply definition, including ply sequence, ply angle, and material properties. While this method considers ply information to some extent, its modeling approach cannot accurately reflect the actual design intent and manufacturing information, such as the specific ply termination and overlap relationships of each beam or rib. More significantly, when modifications to the ply of a specific local structure, such as a skin region connected to a rib, are required during design iterations, engineers must manually locate and modify all relevant region definitions within the complex model. This operation is not only extremely tedious and prone to omissions, leading to model inconsistencies, but also severely restricts the overall efficiency of the design-analysis-optimization iteration process, becoming a major bottleneck in the development process of all-composite aircraft structures.

[0005] In summary, existing composite material aircraft structure modeling methods struggle to balance analytical accuracy with modeling / modification efficiency, failing to meet the urgent demands of modern aircraft, especially rapidly iterating new aircraft, for high-precision and high-efficiency digital design analysis. Therefore, a novel composite material structure modeling method is urgently needed that can accurately reflect design and manufacturing information and support rapid global modifications and iterations. Summary of the Invention

[0006] The technical problem to be solved by the present invention is: in order to overcome the above-mentioned technical problems, the present invention provides a method and equipment for modeling composite material aircraft structures.

[0007] The technical solution adopted by this invention to solve its technical problem is: a composite material aircraft structure modeling method, comprising the following steps: S1. Construct the two-dimensional geometric topology of each component of the aircraft, and construct the shell unit of the component based on the two-dimensional geometric topology; S2. Set the material orientation of the shell unit; S3. Mesh the component to construct a component shell unit mesh; S4. Create a unit set for the component; S5. Construct a single-layer constitutive model of the composite material; S6. Set the unit normal topology of the component shell unit according to the relative positional relationship of each component structure; that is, based on the assembly relationship of the aircraft structure, define the unit normal of each component shell unit in a unified manner according to the preset spatial orientation rules to form the unit normal topology.

[0008] S7. Set up a plywood material model corresponding to the single-layer material constitutive model. The plywood material model includes at least the ply angle and ply thickness, and is associated with the corresponding unit set. S8. Construct the layup sequence topology of the component according to the unit normal topology relationship of the component shell unit; S9. Construct the layup sequence of all components of the entire machine according to the layup sequence topology of the components, and generate a standard format layup data table for the entire machine; the layup data table for the entire machine shall at least include the layup name, layup sequence, material name, material thickness, layup angle of single material layer, and identifier of the associated unit set. S10. Input the entire machine layup data table into computer-aided engineering software, and the software reads the data in the data table to construct the entire machine composite material structure layup model.

[0009] In step S4, the element set refers to the set of shell element meshes for a component. Each element set uniquely corresponds to the structure of each aircraft component; for example, the upper wing skin element set corresponds to the upper wing skin component, and the lower wing skin element set corresponds to the lower wing skin component, etc.

[0010] In step S6, the definition of the unit normal topology follows the spatial orientation principle of front to back, inside to outside, and left to right.

[0011] In step S8, the layering sequence of the component is as follows: from the negative normal of the unit to the positive normal of the unit, and the first layer is counted starting from the negative normal.

[0012] The full-machine layer data table generated in step S9 is an Excel spreadsheet.

[0013] The full-machine layup data table also includes auxiliary fields for output control, manufacturability thickness marking, or layup correction.

[0014] In step S10, the computer-aided engineering software is HyperMesh software, and the whole machine layup data table is imported through the interface or script function of the software.

[0015] The composite material aircraft structure includes a wing, fuselage, horizontal stabilizer, and vertical stabilizer; wherein the wing, horizontal stabilizer, and vertical stabilizer are skin-beam-rib structures, and the fuselage is a skin-frame-beam structure; each component of the aircraft includes at least one of skin, beam web, beam flange, frame web, frame flange, rib web, and rib flange.

[0016] When the composite material aircraft structure is a wing, the layup sequence topology constructed in step S8, in order from the inside out, includes at least the following: wing rib leading edge strip, wing rib middle edge strip, wing rib trailing edge strip, wing front sparsity edge strip, wing rear sparsity edge strip, wing upper skin, and wing lower skin.

[0017] The modeling method further includes step S11: when it is necessary to modify a specific structural layup, locate and modify the data segment corresponding to the specific structure in the whole machine layup data table, and execute step S10 again to update the whole machine composite material structural layup model.

[0018] The present invention provides an electronic device comprising: a memory and a processor, wherein the memory stores a computer program that can run on the processor, and the electronic device executes the program to implement the composite material aircraft structure modeling method of the present invention.

[0019] The present invention provides a computer-readable storage medium storing a computer program thereon, characterized in that the computer program, when executed by a processor, implements the composite material aircraft structure modeling method of the present invention.

[0020] Compared with existing technologies, the composite material aircraft structure modeling method and equipment provided by this invention have the following significant advantages: 1. This invention abandons the crude approach of simply equating composite materials to isotropic materials. Instead, it strictly constructs a material model according to the anisotropic constitutive relationship of a single-layer composite plate and defines the layup based on this. Through the element normal topology and layup sequence topology defined in steps S6 and S8, it ensures that the orientation and stacking order of each layup in three-dimensional space are completely consistent with the design intent. This allows the constructed finite element model to realistically reflect the anisotropic mechanical behavior of the composite structure and the unique coupling effect of the laminate, thus providing a high-precision calculation basis for subsequent strength, stiffness, stability, and dynamic analysis, significantly improving the reliability of the analysis results.

[0021] 2. This invention generates a standardized full-machine layup data table through step S9, structuring and digitizing all modeling information, including complex full-machine geometry, mesh, materials, and layups. When layup design modifications are needed for specific components or local structures, such as a section of beam edge strip or a skin area connected to a specific rib, engineers no longer need to search and edit hundreds or even thousands of element or region attributes one by one in the complex graphical interface of CAE software. They only need to quickly locate the data segment corresponding to the structure in the data table and make centralized modifications. After re-importing into the software, the full-machine model will be automatically updated. This method reduces manual modification work that might have taken hours or even days to a few minutes of data editing, improving modeling and modification efficiency by more than an order of magnitude.

[0022] 3. The model constructed in this invention deeply integrates design and manufacturing information. First, by creating unit sets representing specific components and associating them with these sets through identifier fields such as ESID in the ply data table, each ply in the model can be directly traced back to a specific physical structure. Second, the clearly defined material orientation and unit normal in steps S2 and S6, as well as the ply sequence rule from negative normal to positive normal in step S3, strictly adhere to the actual manufacturing ply sequence and process benchmarks. This allows the analysis model not only to be used for performance calculations but also as a digital basis for verifying manufacturing process data and assessing manufacturability, thus bridging the data chain from design to manufacturing.

[0023] 4. During the multi-round optimization design process of aircraft structures, numerous "design-analysis-modification" iterations are required. This invention transforms the core modeling operation from a graphical interface to data table management, making the updating of the entire aircraft model rapid, controllable, and less prone to errors. Analysis engineers can focus on adjusting design parameters and evaluating results without being bogged down in tedious model reconstruction work. This fundamentally solves the bottleneck problem of traditional methods severely restricting the efficiency of analytical calculation iterations, and is particularly suitable for aircraft models requiring rapid iterative development, such as low-altitude eVTOL aircraft and general aviation aircraft.

[0024] 5. This invention proposes a complete and standardized modeling process and data format. The generated standardized layup data table is independent of specific CAE software sessions, facilitating version management, archiving, and team collaboration. This method is not only applicable to typical components such as wings, fuselages, horizontal stabilizers, and vertical stabilizers, but its core concepts based on topology and data tables can also be extended to the modeling of other complex composite material assembly structures. Furthermore, this data table can be easily integrated with product data management systems, manufacturing execution systems, etc., demonstrating good scalability and promising prospects for digital applications.

[0025] In summary, this invention effectively solves the problem of existing composite material aircraft structure modeling methods being unable to balance accuracy and efficiency, and provides a digital modeling solution that is highly accurate, efficient, traceable, and supports rapid iteration, which is of great value for improving the R&D efficiency of advanced composite material aircraft. Attached Figure Description

[0026] Figure 1 This is a flowchart of the overall process for modeling composite material aircraft structures provided by the present invention.

[0027] Figure 2 This is a schematic diagram of a typical composite material aircraft structure.

[0028] Figure 3 This is an exploded view of a typical box-section structure of an aircraft wing.

[0029] Figure 4 This is a detailed schematic diagram of the wing skin structure.

[0030] Figure 5 These are schematic diagrams of the wing spars structure, where (a) is a schematic diagram of the front wing spars structure and (b) is a schematic diagram of the rear wing spars structure.

[0031] Figure 6 This is a schematic diagram of the wing rib structure.

[0032] Figure 7 This is a schematic diagram of the wing skin-beam-rib adhesive joint structure.

[0033] Figure 8 This is a schematic diagram of the normal topology definition of the wing structure shell unit.

[0034] Figure 9 This is a schematic diagram showing the relationship between the component layup sequence and the unit normal.

[0035] Figure 10 This is a schematic diagram of the topology of the wing skin-beam-rib structure.

[0036] Figure 11 This is a schematic diagram of a typical fuselage box section.

[0037] Figure 12 This is a schematic diagram of the segmented structure of the fuselage beam.

[0038] Figure 13 This is a schematic diagram of the fuselage beam structure.

[0039] Figure 14 This is a schematic diagram of the fuselage frame layout.

[0040] Figure 15 This is a schematic diagram of the fuselage frame structure.

[0041] Figure 16 This is a schematic diagram of the adhesive joint structure of the fuselage skin-frame-beam.

[0042] Figure 17 This is a schematic diagram of the normal topology definition of the fuselage structure shell unit. Among them, (a) is a schematic diagram of the normal topology definition of the fuselage structure shell unit from the front view, and (b) is a schematic diagram of the normal topology definition of the fuselage structure shell unit from the top view.

[0043] Figure 18 This is a topological diagram of the layering sequence of the fuselage skin-frame-beam structure.

[0044] Figure 19 This is an example diagram representing the entire machine's layer-by-layer data.

[0045] Figure 20 This is a schematic diagram of the full-machine composite material structure layup model generated by importing layup data tables into CAE software.

[0046] In the diagram: 100-wing, 101-fuselage, 102-horizontal stabilizer, 103-vertical stabilizer, 1-wing skin, 2-wing rib, 3-wing front spars, 4-wing rear spars, 11-upper wing skin, 12-lower wing skin, 13-wing tail edge, 21-wing rib leading edge slat, 22-wing rib leading web, 23-wing rib center edge slat, 24-wing rib center web, 25-wing rib trailing edge slat, 26-wing 31-Wing front spars upper edge strip, 32-Wing front spars web, 33-Wing front spars lower edge strip, 41-Wing rear spars upper edge strip, 42-Wing rear spars web, 43-Wing rear spars lower edge strip, 5-Fuselage skin, 6-Fuselage right spars, 7-Fuselage left spars, 8-Fuselage frame, 601-Fuselage right spars section 1, 602-Fuselage right spars section 02, 6N-Fuselage right spars section N, 701-Fuselage left spars section 1 702 - Left fuselage beam section 2, 7N - Left fuselage beam section N, 6011 - Right fuselage beam section 1 flange, 6012 - Right fuselage beam section 1 web, 6021 - Right fuselage beam section 2 flange, 6022 - Right fuselage beam section 1 web, 6N1 - Right fuselage beam section N flange, 6N2 - Right fuselage beam section 1 web, 7011 - Left fuselage beam section 1 flange, 7012 - Left fuselage beam section 1 web, 7021 - Left fuselage beam section 2 Flanging strip, 7022-Fuselage left beam section 2 web plate, 7N1-Fuselage left beam section N flange strip, 7N2-Fuselage left beam section N web plate, 801-Fuselage frame 1, 802-Fuselage frame 2, 8N-Fuselage frame N, 8011-Fuselage frame 1 flange strip, 8012-Fuselage frame 1 web plate, 8021-Fuselage frame 2 flange strip, 8022-Fuselage frame 2 web plate, 8N1-Fuselage frame N flange strip, 8N2-Fuselage frame N web plate. Detailed Implementation

[0047] The invention will now be described in further detail with reference to the accompanying drawings. It should be emphasized that the following description is merely exemplary and not intended to limit the scope or application of the invention.

[0048] like Figure 1As shown, this invention provides a method for modeling composite material aircraft structures. Its core lies in using a layup composite material modeling method as a foundation, then rationally defining the component element normal topology and component layup sequence topology, constructing standard EXCEL data tables for each component layup, and then combining these EXCEL data tables into an EXCEL data table for the entire aircraft composite material structure layup. Finally, the entire aircraft composite material structure layup model data EXCEL table is input into software (HyperMesh software, OptiStruct solver) to automatically construct the entire aircraft composite material structure layup. This modeling method can reflect the overall aircraft structural design and manufacturing information. When modifications are needed for specific structural layups, component layup information can be quickly retrieved from the layup model data EXCEL table, and only the specific structural layup data segment needs to be modified to update the overall aircraft layup information, thereby updating the entire aircraft composite material structure layup. This improves the efficiency of overall aircraft modeling and modification, as well as the efficiency of analysis and calculation iterations.

[0049] The following explanation uses a typical composite material aircraft structure as an example. Figure 2 As shown, it typically includes a wing 100, a fuselage 101, a horizontal stabilizer 102, and a vertical stabilizer 103. The wing 100, horizontal stabilizer 102, and vertical stabilizer 103 are skin-beam-rib structures, while the fuselage 101 is a skin-frame-beam structure.

[0050] Example 1: Wing Modeling Based on Skin-Beam-Rib Structure

[0051] This embodiment uses an aircraft wing as an example, whose typical structure is a box segment composed of skin, beams, and ribs. For example... Figure 3 , 7 As shown, the composite material structure aircraft wing consists of wing skin 1, wing front spars 3, wing rear spars 4, and wing ribs 2. Figure 4 As shown, the wing skin consists of the upper wing skin 11, the lower wing skin 12, and the wing trailing edge 13. Figure 5 As shown, the wing's front sparsity consists of a front sparsity web 32, a front sparsity upper edge strip 31, and a front sparsity upper edge strip 33. The wing's rear sparsity consists of a rear sparsity web 42, a rear sparsity upper edge strip 41, and a rear sparsity upper edge strip 43. Figure 6 As shown, the wing rib consists of the leading edge strip 21, the leading web 22, the center edge strip 23, the center web 24, the trailing edge strip 25, and the trailing web 26. The wing skin, spars, and ribs are joined using adhesive bonding; details of the adhesive bonding joint structure can be found in [link to details]. Figure 7 At the front spar joint, the bonding structure is as follows: wing rib center edge strip – wing front spar edge strip, web plate – wing leading edge strip – upper and lower wing skin bonding. At the rear spar joint, the bonding structure is as follows: wing rib center edge strip – wing rear spar edge strip, web plate – wing rib rear edge strip – upper and lower wing skin bonding.

[0052] Step 1: Component geometry and topology and shell element mesh

[0053] Based on the mid-surface (or design reference surface) of the component, construct its two-dimensional (2D) geometric topology.

[0054] Based on this 2D geometric topology, shell elements are constructed in CAE software such as HyperMesh to characterize the component, and the element properties are set to composite material properties, such as PCOMPP in the OptiStruct solver.

[0055] Mesh the shell elements to generate a high-quality shell element mesh for the component. For composite adhesive-bonded assembly joints, such as the junction of skin and beams / ribs, use common-node simulation and merge the adhesive surfaces into a single surface during geometry processing. This step corresponds to... Figure 3 and Figure 7 The structural relationship between the wing and the wing joint is shown.

[0056] Step 2: Define the element set and material properties

[0057] For efficient allocation of subsequent ply properties, element sets (SETs) need to be created for different structural components such as skin, beam web, beam flange, frame web, frame flange, rib web, rib flange, and trailing edge. An element set refers to a collection of shell element meshes for a component. Each element set uniquely corresponds to a structural component of the aircraft; for example, the upper wing skin element set corresponds to the upper wing skin component, and the lower wing skin element set corresponds to the lower wing skin component, etc.

[0058] Using the TCL / TK language, cell sets are created in batches within the software, with ID numbers assigned as 1, 2, 3, ... These ID numbers serve as the identifiers for the cell sets. The following explanation uses the TCL language as an example: Using the TCL software language, write a TCL script file to create cell sets. First, create an empty cell set in the script file, specifying its name and ID number. If no ID number is specified, the software defaults to the order in which cell sets are created, with ID numbers starting from 1 and increasing sequentially. The ID format is a positive integer, i.e., 1, 2, 3… The ID number is the identifier of the cell set.

[0059] The software loads the TCL script file, creating a batch of element sets. At this stage, the created element sets do not yet contain shell elements associated with components. The software then selects the corresponding components to form shell elements for each element set and associates these shell elements with the element set.

[0060] Each unit set should uniquely correspond to a component's structural composition. The naming convention is to follow the hierarchical relationship of the component structure, reflecting the corresponding component's structural composition and ensuring consistency with its name. The format is XX_XX_XX…, for example, wing_front sparsa_upper edge strip; wing_front sparsa_web; wing_wing rib_leading edge strip; wing_tail edge; wing_upper skin.

[0061] The TCL language name format for creating a unit set is as follows. You only need to change the unit set name and ID number to create unit sets for other components.

[0062] createentity sets cardimage="SET_ELEM" name="Wing_Rib_Front_YT" id=1 createentity sets cardimage="SET_ELEM" name="Wing_Rib_Front_FB" id=2 createentity sets cardimage="SET_ELEM" name="Wing_FrontBeam_FB" id=3 The following table shows examples of SET (set) naming conventions (English letters are generally used in software):

[0063] Construct a constitutive model of a single-layer composite material. Under plane stress, the stress-strain relationships of the single-layer orthotropic material in the principal directions are as follows: E1, E2, ν 12 G 12 These constitute the input parameters for the single-layer material model.

[0064] ; ; In the formula: E1: Elastic modulus of the material in the principal direction 1; E2: Biaxial elastic modulus of the material in the principal directions; ν 12 ,ν 21 Poisson's ratio in the plane; G 12 In-plane shear modulus; Stress in the principal direction 1 of the material; : Stress in two directions along the principal directions of the material; Shear stress in planes 1 and 2 along the principal directions of the material; Strain in the principal direction 1 of the material; Material strain in two directions (principal and secondary). Shear strain in planes 1 and 2 along the principal directions of the material.

[0065] Step 3: Define the element normal topology and ply sequence topology

[0066] This step forms the basis for building standardized data tables, and its rules ensure consistency in the design and analysis models.

[0067] Define the normal topology of the unit: Based on the actual assembly and stacking sequence of the aircraft structure, the positive normal directions of all component shell units are uniformly defined. For the wing, according to the relative positional relationship of the wing skin-spar-rib structure, and referring to the principles of front to back, inside to outside, and left to right, the normal direction topology of the wing structural shell units is defined. See details. Figure 8 : The normal direction of the upper skin unit is defined as upward (pointing to the outer side of the upper skin of the wing).

[0068] The normal direction of the lower skin unit is defined as downward (pointing to the outer side of the lower skin of the wing).

[0069] The normal direction of the beam web element is defined as pointing from the front beam to the rear beam (i.e., the chord direction of the wing).

[0070] The normal direction of the rib and web unit is defined as pointing from the left and right symmetrical planes of the aircraft to the wingtip (i.e., the spanwise direction).

[0071] The positive normal of the flange elements of beams and ribs specifies that the normal of the connected structural component elements should be consistent. The flanges and their corresponding connected structures are modeled using a shared-node approach.

[0072] Establish the layer sequence topology: Figure 9 As shown, the layering sequence of the components is as follows: from the negative normal of the element to the positive normal of the element, the first layer starts from the negative normal. Figure 9 In this diagram, 1-8 represent a total of 8 layers, with the first layer starting from the negative normal direction of the unit. Based on this rule and the relative positions of the components, the layer sequence topology of the entire machine can be derived. For example, Figure 10 As shown, for the wing skin-spar-rib structure, the layup sequence (from inside to outside) can be: wing rib leading edge strip → wing rib leading web → ... → wing trailing sparsity strip → upper / lower wing skin → wing trailing edge.

[0073] The layup sequence must accurately reflect the relative positional relationships of the wing structural components. The layup sequence starts from the negative normal direction of the element. The innermost / foremost / leftmost component corresponding to the negative normal direction in the element normal topology should be laid up first. At the same time, the relative positional relationships of the components must be considered, and the structural components should be defined in sequence from front to back, from inside to outside, and from left to right.

[0074] The relative positions of the wing structural components are as follows: (1) Relative positional relationship from inside to outside: ① Wing rib leading edge strip - upper wing skin, lower wing skin;

[0075] ③Wing trailing edge strip - upper wing skin, lower wing skin.

[0076] (2) Relative positional relationship from front to back: Wing rib leading edge strip - wing front spars web - wing rib middle edge strip - wing rear spars web - wing rib trailing edge strip - wing trailing edge.

[0077] (3) Relative positions from left to right: front web of wing rib, middle web of wing rib, rear web of wing rib.

[0078] Therefore, the wing structure layup sequence can be given according to the unit normal topology and the relative positional relationship of the component structures: ① From front to back and from left to right, complete the layup of the leading edge strip of the wing rib and the front web of the wing rib; ②From front to back, from left to right, the web of the wing front spars, the center edge strip of the wing ribs, the center web of the wing ribs, and the web of the wing rear spars are laid up. ③ Following the order from front to back and from left to right, complete the plying of the trailing edge strip of the wing rib, the rear web of the wing rib, and the tail edge of the wing; ④ Following the process from the inside out, complete the layup of the upper and lower edge strips of the wing front spars and the upper and lower edge strips of the wing rear spars; ⑤ Complete the upper and lower skin layers of the wing from the inside out.

[0079] Step 4: Generate wing ply data table

[0080] Generate a data table: Based on the above ply sequence topology, associate the information of each ply (Ply Name, Ply ID, Mat Name, Thickness, Orientation) with its corresponding component unit set ESID, and fill it into a standard format table, such as an Excel spreadsheet. A standard format Excel spreadsheet refers to an Excel spreadsheet that meets the software requirements. An Excel spreadsheet conforming to the HyperMesh software requirements has 9 columns of information, and the content of each column is required as follows: Column 1: Ply Name; Column 2: Sequence Number of Single-Layer Ply in the Overall Machine Ply Structure; Column 3: Material Name; Column 4: Material Thickness; Column 5: Single-Layer Material Ply Angle; Column 6: Single-Layer Material Calculation Results Output; Column 7: Single-Layer Material Manufacturing Thickness; Column 8: Ply Angle and Thickness Correction; Column 9: Unit Set ID Number. Columns 7 and 8 are not mandatory and can be empty, but they cannot be deleted. Column 6 is usually required; if empty, the single-layer ply calculation results will not be output. All other columns are mandatory.

[0081] Example 2: Fuselage Modeling Based on Skin-Frame-Beam Structure

[0082] This embodiment focuses on an aircraft fuselage, whose typical structure consists of skin, frame, and beams. Composite material aircraft fuselages generally consist of fuselage skin, fuselage frame, left fuselage beam, and right fuselage beam; see details below. Figure 11 Generally, the left and right fuselage beams can be divided into multiple sections from the nose to the tail, as detailed below. Figure 12 For the fuselage frame, multiple frames can be arranged from the nose to the tail; see details. Figure 14 The fuselage skin consists of the left fuselage skin and the right fuselage skin; the fuselage beams (left fuselage beam and right fuselage beam) consist of beam flanges and beam webs, see details. Figure 13 The fuselage frame consists of frame flanges and frame webs, see details. Figure 15 The fuselage skin, frame, and beams utilize an adhesive bonding structure. The adhesive joint structure is as follows: fuselage beam flange—fuselage frame flange, web—fuselage skin adhesive bonding, or fuselage beam flange—fuselage frame flange, web adhesive bonding. See details... Figure 16 .

[0083] Step 1: Component geometry and topology and shell element mesh

[0084] Based on the mid-surface (or design reference surface) of the component, construct its two-dimensional (2D) geometric topology.

[0085] Based on this 2D geometric topology, shell elements are constructed in CAE software such as HyperMesh to characterize the component, and the element properties are set to composite material properties, such as PCOMPP in the OptiStruct solver.

[0086] Mesh the shell elements to generate a high-quality shell element mesh for the component. For composite adhesive-bonded assembly joints, such as the junction of skin and beams / ribs, use common-node simulation and merge the adhesive surfaces into a single surface during geometry processing. This step corresponds to... Figure 11 and Figure 16 The structural relationship between the fuselage and fuselage joints is shown.

[0087] Step 2: Define the element set and material properties

[0088] For efficient allocation of subsequent layup properties, unit sets (SETs) need to be created for different fuselage components.

[0089] Using the TCL software language, write a TCL script file to create cell sets. First, create an empty cell set in the script file, specifying its name and ID number. If no ID number is specified, the software defaults to the order in which cell sets are created, with ID numbers starting from 1 and increasing sequentially. The ID format is a positive integer, i.e., 1, 2, 3… The ID number is the identifier of the cell set.

[0090] The software loads the TCL script file, creating a batch of element sets. At this stage, the created element sets do not yet contain shell elements associated with components. The software then selects the corresponding components to form shell elements for each element set and associates these shell elements with the element set.

[0091] Each unit set should uniquely correspond to a component's structural composition. The naming convention is to follow the hierarchical relationship of the component structure, reflecting the corresponding component structure and ensuring consistency with its name. The format is XX_XX_XX…. For example, Fuselage_Frame1_Flange; Fuselage_Frame1_Body Plate; Fuselage_Right Beam 1_Flange; Fuselage_Left Skin, etc.

[0092] The TCL language name format for creating a unit set is as follows. You only need to change the unit set name and ID number to create unit sets for other components.

[0093] createentity sets cardimage="SET_ELEM" name="Fuselage_Frame1_FB" id=14

[0094] Construct a constitutive model of a single-layer composite material. Under plane stress, the stress-strain relationships of the single-layer orthotropic material in the principal directions are as follows: E1, E2, ν 12 G 12 These constitute the input parameters for the single-layer material model.

[0095] ; ; In the formula: E1: Elastic modulus of the material in the principal direction 1; E2: Biaxial elastic modulus of the material in the principal directions; ν 12 ,ν 21 Poisson's ratio in the plane; G 12 In-plane shear modulus; Stress in the principal direction 1 of the material; : Stress in two directions along the principal directions of the material; Shear stress in planes 1 and 2 along the principal directions of the material; Strain in the principal direction 1 of the material; Material strain in two directions (principal and secondary). Shear strain in planes 1 and 2 along the principal directions of the material.

[0096] Step 3: Define the element normal topology and ply sequence topology

[0097] Define the element normal topology: According to the relative positional relationship of the fuselage skin-frame-beam structure, and referring to the principle of front to back, inside to outside, and left to right, define the normal direction topology of the fuselage structure shell element, as shown in Figure 17.

[0098] The layup order points from the negative normal of the element to the positive normal of the element. Therefore, according to the relative positional relationship of the components, from front to back, from inside to outside, and from left to right, the frontmost / inner / leftmost side of the component is defined as the negative normal direction of the element.

[0099] The normal direction of the left fuselage skin unit is defined as pointing from the left wing root to the wingtip; the normal direction of the right fuselage skin unit is defined as pointing from the right wing root to the wingtip; the normal direction of the fuselage frame web unit is defined as pointing from the nose to the tail; the normal direction of the left fuselage beam web unit is defined as pointing from the left wing root to the wingtip; the normal direction of the right fuselage beam web unit is defined as pointing from the right wing root to the wingtip; the normal direction of the fuselage frame flange and fuselage beam flange units is consistent with the normal direction of the connecting structural component units. Because the flange and its corresponding connecting structure are modeled using a shared node method, the normal direction of the fuselage frame flange and fuselage beam flange units not connected to other structures is defined as pointing from the inner surface of the flange to the outer surface of the flange. Establish the layup sequence topology: Figure 9 As shown, the layering sequence of the components is as follows: from the negative normal of the element to the positive normal of the element, the first layer starts from the negative normal. Figure 9 In this context, 1-8 represent a total of 8 layers, with the first layer starting from the negative normal direction of the unit. Based on this rule and the relative positions of the components, the layup sequence topology of the entire aircraft can be derived. For the fuselage skin-frame-beam structure, Figure 18 As shown, the layup sequence (from inside to outside) may be: fuselage 01 frame web plate → fuselage left / right beam 01 segment edge strip → ... → fuselage N frame edge strip → fuselage left / right skin.

[0100] The layup sequence must accurately reflect the relative positional relationships of the fuselage structural components. The layup sequence starts from the negative normal direction of the unit. The innermost / foremost / leftmost components corresponding to the negative normal direction in the unit normal topology should be laid up first. At the same time, the relative positional relationships of the components must be considered, and the structural components should be defined in sequence from front to back, from inside to outside, and from left to right.

[0101] The relative positions of the fuselage structural components are as follows: (1) Relative positional relationship from inside to outside:

[0102] (2) Relative positional relationship from front to back: Fuselage frame 1 web plate - fuselage left beam / right beam 1 section flange - fuselage frame 2 web plate - fuselage left beam / right beam 2 section flange - ... - fuselage frame N web plate.

[0103] (3) From left to right: left fuselage beam web, right fuselage beam web.

[0104] Therefore, the wing structure layup sequence can be given according to the unit normal topology and the relative positional relationship of the component structures: Step 1: From front to back and from left to right, complete the layup of the fuselage frame 1 web plate, fuselage left / right beam 1 section flange strip, and fuselage left / right beam 1 section web plate components; Step 2: From front to back and from left to right, complete the layup of the fuselage frame 2 web plates, fuselage left beam / right beam 2-section flange strips, and fuselage left beam / right beam 2-section web plate components; … Step N: Complete the web layering of the Nth frame of the fuselage from front to back; Step N+1: Complete the layup of the fuselage frame edge strip from the inside out; … Step 2N: From the inside out, complete the N-layer layup of the fuselage frame edge strips; Step 2N+1: Complete the skin layering of the left fuselage from the inside out; Step 2N+2: Complete the skin layup on the right side of the fuselage from the inside out.

[0105] Step 4: Generate fuselage layup data table

[0106] Generate a data table: Based on the above ply sequence topology, associate the information of each ply (Ply Name, Ply ID, Material Name, Thickness, Orientation) with its corresponding component element set ESID, and fill it into a standard format table, such as an Excel spreadsheet. The Excel spreadsheet required by the HyperMesh software should have 9 columns, with each column containing the following information: Column 1: Ply Name; Column 2: Sequence Number of Single-Layer Ply in the Overall Machine Ply Structure; Column 3: Material Name; Column 4: Material Thickness; Column 5: Single-Layer Material Ply Angle; Column 6: Single-Layer Material Calculation Results Output; Column 7: Single-Layer Material Manufacturing Thickness; Column 8: Ply Angle and Thickness Correction; Column 9: Unit Set ID Number. Columns 7 and 8 are not mandatory and can be empty, but they cannot be deleted. Column 6 is usually required; if empty, the single-layer ply calculation results will not be output. All other columns are mandatory.

[0107] Full-machine model integration and construction: I. Integrating the Overall Aircraft Data Sheets: The horizontal and vertical stabilizers are wing-like structures, and their plywood data can be constructed using the same method as the wing. The wing plywood data sheets from Example 1, the fuselage plywood data sheets from Example 2, and the plywood data sheets for other components such as the horizontal and vertical stabilizers are combined according to the overall aircraft assembly relationship to form a complete standard Excel sheet for the overall aircraft composite material structure plywood model data, as shown below. Figure 19 As shown.

[0108] Figure 19The Excel spreadsheet only contains a portion of the total layer data for the entire machine. Here, Ply Name represents the layer name, such as Wing_Rib_Front_YT_P1, Wing_Rib_Front_YT_P1... Wing_SkinDown_P6; Ply ID represents the layer order for the entire machine, for example, a total of 300 layers for all components. The corresponding Ply IDs are 1, 2, ..., 300; IDCard image represents the card attribute, where PLY is the ply attribute identifier for each layer of the card in the hypermesh software. Each ply consists of one PLY, and all PLYs are arranged in a specified order to form the ply of the entire machine; Mat Name represents the material name; Thickness represents the thickness of a single layer of material, in mm; Orientation represents the ply angle, in °; Output Results represents the output results, yes indicates that the software outputs the calculation results data of this layer of PLY in the calculation results, such as stress, strain, displacement, etc.; ESID represents the component element set ID sequence number, 1 in the table is the identifier of the component element set, different components correspond to different element set identifiers; TMANU represents the manufacturable thickness of a single layer (default, no need to fill in); DRAPE_ID represents the ply angle and thickness correction ID (generally no need to set, keep the default value of 0). T300 represents carbon fiber fabric, a material name defined in the software, with the following parameters: 1-direction elastic modulus E1=59000MPa, 2-direction elastic modulus E2=58000MPa, in-plane shear modulus G12=2000MPa, and in-plane Poisson's ratio v12=0.03. T700 represents carbon fiber unidirectional tape, with 1-direction elastic modulus E1=127000MPa, 2-direction elastic modulus E2=6700MPa, in-plane shear modulus G12=3300MPa, and in-plane Poisson's ratio v12=0.315.

[0109] In the layup data Excel sheet, manually enter the ID number of the cell set.

[0110] II. Driven Modeling: In CAE software, the entire machine's ply data table is read through built-in interfaces, scripts, or dedicated programs. The software locates the corresponding element set based on the ESID in the ply data table and automatically assigns ply attributes to these elements according to Ply ID order, Orientation angle, and Mat Name, thereby efficiently and accurately constructing a complete finite element model of the entire composite material.

[0111] This embodiment uses HyperMesh's built-in batch import function to import the full-machine composite material structure layup data table, thereby automatically generating the full-machine composite material structure layup model. The implementation method is as follows: In the HyperMesh software's CompositeBrowser browser, perform the following operations: Step 1: Right-click in the blank area and create a Laminate; Step 2: Right-click on Laminate and select Import Spreadsheet; Step 3: In the pop-up window, select "Open Full Machine Layup Table," and the software will automatically generate a full machine composite structure layup model. See [link / reference]. Figure 20 .

[0112] Rapid modification and iteration of the model: When design changes require modifications to specific structures, such as the ply of the upper edge slats of a wing's front spars, it's unnecessary to modify the unit attributes one by one in the CAE software's graphical interface. Instead, simply locate all ply data rows related to that structure in the overall aircraft ply data table by filtering or searching for ESIDs, and make centralized modifications, such as adjusting angles or adding / removing plies. Then, re-execute the driven modeling process in step four to quickly update the entire analysis model, significantly improving design iteration efficiency.

[0113] The two specific embodiments above illustrate in detail the entire application process of the method of the present invention to two typical aircraft structures: wings and fuselages. By defining the topology and then generating data tables to drive modeling, the contradiction between accuracy and efficiency in traditional methods is successfully resolved.

[0114] Based on the above-described preferred embodiments of the present invention, and through the foregoing description, those skilled in the art can make various changes and modifications without departing from the inventive concept. The technical scope of this invention is not limited to the contents of the specification, but must be determined according to the scope of the claims.

Claims

1. A method for modeling composite material aircraft structures, characterized in that, Includes the following steps: S1. Construct the two-dimensional geometric topology of each component of the aircraft, and construct the shell unit of the component based on the two-dimensional geometric topology; S2. Set the material orientation of the shell unit; S3. Mesh the component to construct a component shell unit mesh; S4. Create a unit set for the component; S5. Construct a constitutive model of a single-layer composite material; S6. Set the element normal topology of the shell element according to the relative positional relationship of each component structure; S7. Set up a plywood material model corresponding to the single-layer material constitutive model. The plywood material model includes at least the ply angle and the ply thickness, and is associated with the corresponding unit set. S8. Construct the layup sequence topology of the component according to the unit normal topology relationship of the component shell unit; S9. Construct the layup sequence of all components of the entire machine according to the layup sequence topology of the components, and generate a standard format layup data table for the entire machine; the layup data table for the entire machine shall at least include the layup name, layup sequence, material name, material thickness, layup angle of single material layer, and identifier of the associated unit set. S10. Input the full machine layup data table into computer-aided engineering software, and the software reads the data in the data table to construct a full machine composite material structure layup model.

2. The composite material aircraft structure modeling method as described in claim 1, characterized in that: In step S6, the definition of the unit normal topology follows the spatial orientation principle of front to back, inside to outside, and left to right.

3. The composite material aircraft structure modeling method as described in claim 1, characterized in that: In step S8, the layering sequence of the component is as follows: from the negative normal of the unit to the positive normal of the unit, and the first layer is counted starting from the negative normal.

4. The composite material aircraft structure modeling method as described in claim 1, characterized in that: The full-machine layer data table generated in step S9 is an Excel spreadsheet.

5. The composite material aircraft structure modeling method as described in claim 4, characterized in that: The full-machine layup data table also includes auxiliary fields for output control, manufacturability thickness marking, or layup correction.

6. The composite material aircraft structure modeling method as described in claim 1, characterized in that: In step S10, the computer-aided engineering software is HyperMesh software, and the whole machine layup data table is imported through the interface or script function of the software.

7. The composite material aircraft structure modeling method as described in claim 1, characterized in that: The composite material aircraft structure includes a wing, fuselage, horizontal stabilizer, and vertical stabilizer; wherein the wing, horizontal stabilizer, and vertical stabilizer are skin-beam-rib structures, and the fuselage is a skin-frame-beam structure; each component of the aircraft includes at least one of skin, beam web, beam flange, frame web, frame flange, rib web, and rib flange.

8. The composite material aircraft structure modeling method as described in claim 7, characterized in that: When the composite material aircraft structure is a wing, the layup sequence topology constructed in step S8, in order from the inside out, includes at least the following: wing rib leading edge strip, wing rib middle edge strip, wing rib trailing edge strip, wing front sparsity edge strip, wing rear sparsity edge strip, wing upper skin, and wing lower skin.

9. The composite material aircraft structure modeling method as described in claim 1, characterized in that: The modeling method further includes step S11: when it is necessary to modify a specific structural layup, locate and modify the data segment corresponding to the specific structure in the whole machine layup data table, and execute step S10 again to update the whole machine composite material structural layup model.

10. An electronic device, characterized in that, include: A memory and a processor, the memory storing a computer program executable on the processor, wherein the electronic device, when executing the program, implements the composite material aircraft structure modeling method according to any one of claims 1-9.

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