Finite element modeling method and device for composite material ultra-large aspect ratio UAV

Abstract

The present invention relates to the technical field of unmanned aerial vehicle design, and specifically discloses a finite element modeling method and device for a composite material ultra-large aspect ratio unmanned aerial vehicle, the method comprising: dividing the unmanned aerial vehicle into a plurality of structural component sub-models according to the overall layout of the entire aircraft; determining the structural type of the elements of the structural components of the unmanned aerial vehicle, selecting a sub-model simplification principle and establishing a unit; creating materials, creating and assigning unit attributes based on the created materials; renumbering each sub-model and its nodes / units / materials / attributes according to a numbering principle; debugging each sub-model until it runs smoothly; integrating all sub-models to obtain a finite element model of the entire aircraft, and debugging the model until it runs smoothly. The present invention provides corresponding simplification principles for different structural components of an aircraft, and unifies modeling standards, regulates element numbering, reduces model size, and improves computational efficiency, while ensuring that the basic geometric features, force transmission paths, stiffness, and mass characteristics of the structure are not changed after simplification.

Description

Technical Field

[0001] The present invention relates to the technical field of unmanned aerial vehicle (UAV) design, and in particular to a finite element modeling method and device for a composite material ultra-large aspect ratio UAV. Background Art

[0002] Composite ultra-high-aspect-ratio drones are lightweight and achieve a closed-circuit design weight by maximizing the designability of laminated composite materials. This design is primarily affected by gust loads. When establishing a full-aircraft finite element model, in addition to general modeling principles, a static strength full-aircraft finite element model of the composite ultra-high-aspect-ratio drone must be developed based on its unique structural design characteristics. This model must be suitable for analyzing and calculating the drone's structural strength, stiffness, and stability, and support the design of its dynamic strength, aeroelasticity, durability, and damage tolerance.

[0003] The existing finite element model of the static strength of the entire composite material ultra-large aspect ratio UAV usually has problems such as inconsistent modeling standards for multiple components, irregular element numbering, and irregular model size reduction. These problems lead to poor readability of model parameters, weak operability, and low computational efficiency, which greatly affects the modeling accuracy and efficiency.

[0004] Based on this technical background, the present invention studies a finite element modeling method and device for a composite material ultra-large aspect ratio UAV. Summary of the Invention

[0005] In response to the shortcomings of the existing technology, the present invention provides a finite element modeling method and device for a composite material ultra-large aspect ratio unmanned aerial vehicle. The method provides corresponding simplification principles for different structural components in the aircraft. Under the premise of ensuring that the basic geometric characteristics, force transmission path, stiffness and mass characteristics of the structure are not changed after simplification, the method unifies the modeling standards, regulates the element numbering, reduces the model size, increases the readability of the calculation file and improves the calculation efficiency.

[0006] To achieve the above objectives, a first aspect of the present invention provides a finite element modeling method for a composite material ultra-high aspect ratio UAV, comprising:

[0007] According to the overall layout of the whole aircraft, the UAV is divided into multiple structural component sub-models;

[0008] Determining the structural type of the components of the UAV structural components, selecting the sub-model simplification principle and establishing units;

[0009] Create materials, create and assign unit properties based on the created materials;

[0010] Renumber each sub-model and its nodes / elements / materials / attributes according to the numbering principle;

[0011] Debug each sub-model until it runs smoothly;

[0012] Integrate all sub-models to obtain the finite element model of the entire aircraft, and debug it until it runs smoothly.

[0013] A second aspect of the present invention provides a finite element modeling device for a composite material ultra-large aspect ratio UAV, comprising:

[0014] The sub-model division module is used to divide the UAV into multiple structural component sub-models according to the overall layout of the whole aircraft;

[0015] A unit establishment module, configured to determine the structural type of the components of the UAV structural components, select the sub-model simplification principle and establish a unit;

[0016] Material creation module, used to create materials, create and assign unit properties based on the created materials;

[0017] The renumbering module is used to renumber each sub-model and its nodes / elements / materials / attributes according to the numbering principle;

[0018] Sub-model debugging module, used to debug each sub-model until it runs smoothly;

[0019] The finite element model debugging module is used to integrate all sub-models to obtain the finite element model of the entire aircraft, and debug it until it runs smoothly.

[0020] A third aspect of the present invention provides an electronic device, comprising:

[0021] a memory storing executable instructions;

[0022] A processor runs the executable instructions in the memory to implement the finite element modeling method for the composite material ultra-large aspect ratio UAV according to the first aspect.

[0023] A fourth aspect of the present invention provides a computer-readable storage medium storing a computer program, which, when executed by a processor, implements the finite element modeling method for the composite material ultra-large aspect ratio UAV described in the first aspect.

[0024] The beneficial effects of the present invention include:

[0025] (1) The finite element modeling method for composite material ultra-large aspect ratio UAV proposed in this invention provides corresponding simplification principles for different structural components in the aircraft. Under the premise of ensuring that the basic geometric characteristics, force transmission path, stiffness and mass characteristics of the structure are not changed after simplification, the modeling standards are unified, the element numbering is standardized, the model size is reduced, the calculation file readability is increased and the calculation efficiency is improved.

[0026] (2) The finite element modeling method for composite material ultra-large aspect ratio UAV proposed in the present invention unifies the modeling standards: multiple components are created simultaneously with reference to the same modeling standards, which improves the modeling efficiency and realizes the non-differentiation of modeling by multiple people; the elements are numbered in a regular manner, and the nodes / units / materials / attributes of the input files submitted for calculation are numbered in a regular and orderly manner, with good readability and strong operability.

[0027] (3) The finite element modeling method for composite material ultra-large aspect ratio UAV proposed in this invention reduces the model size and adopts the simplified principle of one-dimensional / two-dimensional unit composite to improve the calculation efficiency without changing the basic geometric characteristics, force transmission path, stiffness and mass characteristics of the structure.

[0028] Other features and advantages of the present invention will be described in detail in the following detailed description. BRIEF DESCRIPTION OF THE DRAWINGS

[0029] The above and other objects, features and advantages of the present invention will become more apparent through a more detailed description of exemplary embodiments of the present invention with reference to the accompanying drawings.

[0030] Figure 1 This is a flow chart of the finite element modeling method for the composite material ultra-large aspect ratio UAV proposed in the present invention.

[0031] Figure 2 This is a flow chart of a specific implementation of the finite element modeling method for a composite material ultra-large aspect ratio UAV proposed in the present invention.

[0032] Figure 3 This is a simplified schematic diagram of a thin-walled circular tube beam unit in a specific implementation of the finite element modeling method for a composite material ultra-large aspect ratio UAV proposed in the present invention. DETAILED DESCRIPTION

[0033] The preferred embodiments of the present invention will be described in more detail below. Although the preferred embodiments of the present invention are described below, it should be understood that the present invention can be implemented in various forms and should not be limited to the embodiments set forth herein.

[0034] The present invention provides a finite element modeling method for a composite material ultra-large aspect ratio UAV, such as Figure 1 Shown, including:

[0035] According to the overall layout of the whole aircraft, the UAV is divided into multiple structural component sub-models;

[0036] Determine the structural type of the components in the UAV structure, select the sub-model simplification principle and establish the unit;

[0037] Create materials, create and assign unit properties based on the created materials;

[0038] Renumber each sub-model and its nodes / elements / materials / attributes according to the numbering principle;

[0039] Debug each sub-model until it runs smoothly;

[0040] Integrate all sub-models to obtain the finite element model of the entire aircraft, and debug it until it runs smoothly.

[0041] The method of the present invention provides corresponding simplification principles for different structural components in an aircraft. On the premise of ensuring that the basic geometric characteristics, force transmission paths, stiffness and mass characteristics of the structure are not changed after simplification, it unifies the modeling standards, regulates the element numbering, reduces the model size, increases the readability of the calculation file and improves the calculation efficiency.

[0042] According to the present invention, a plurality of structural component sub-models are obtained by dividing the UAV into components;

[0043] UAVs are divided into wing sections, fuselage, tail and landing gear according to their components;

[0044] The structural components in all sub-models follow the same simplification principle to achieve standardized and non-differentiated modeling among the sub-models.

[0045] According to the present invention, determining the structural type of the components of the UAV structural parts, selecting the sub-model simplification principle and establishing the unit includes:

[0046] If the structural form of the component elements is simple and can be loaded by tension or compression, it can be simplified to shell elements;

[0047] If the structural form of the constituent elements is irregular and the model has multi-dimensional combinations of surfaces and lines after geometric pre-processing, a composite simplification using shell elements and beam elements is used;

[0048] If the structural form of the component is simple and can only bear tension, it can be simplified into a membrane unit;

[0049] If the component is a connection joint of a non-primary load-bearing component, it can be simplified to an RBE2 rigid constraint unit.

[0050] In the present invention, a unified modeling standard is adopted: multiple components are created simultaneously with reference to the same modeling standard, which improves modeling efficiency and enables non-differentiated modeling by multiple people; element numbering is regular, and the node / unit / material / attribute numbering of the input file submitted for calculation is regular and orderly, with good readability and strong operability.

[0051] According to the present invention, creating a material includes:

[0052] Determine the material type, determine whether to simplify the material based on the material type, and simplify the material that needs to be simplified;

[0053] Determine whether to simplify the material based on the material type. The simplification of materials that need to be simplified includes:

[0054] If it is an isotropic material or an orthotropic material, no simplification is done and its performance parameters are determined based on the performance test data;

[0055] If it is a laminated composite material, it is simplified to an isotropic material. The simplification principle is as follows: the comprehensive performance of the laminated composite material is calculated using finite element software, and the isotropic material is created using the equivalent elastic modulus of the beam element in the direction of the axial force;

[0056] If there are multiple layup methods, they can be simplified into a unified laminated composite material. The simplification principle is: based on the proportion of all layup methods in the region, the layup method with the largest coverage area is used to create a regional laminated composite material.

[0057] According to the present invention, creating and assigning unit attributes based on the created material includes:

[0058] Define the cross-sectional size information of the beam element according to the actual cross-sectional shape, and assign the cross-sectional information and material information to the beam element;

[0059] Assign material information to the corresponding shell / membrane elements according to the actual structure;

[0060] According to the actual ply situation, in the direction where the ply change affects the structural performance relatively obviously, Nc = 5 to 8 attribute zones are divided. In the direction perpendicular to it, Nk = 1 / 10 to 1 / 8 times the aspect ratio attribute zones are divided. That is, within this range, the model is divided into Nc*Nk attribute zones in total.

[0061] Preferably, the numbering principle is: use the sub-model as the first identifier of the number to avoid repeated numbering of nodes / units / materials between different sub-models.

[0062] According to the present invention, renumbering each sub-model and its nodes / elements / materials / attributes includes:

[0063] The file name of the sub-model reflects the sub-model number, design round number and modeling time;

[0064] The node number is 8 digits. The first three digits represent the structural component submodel, the structural component's location in the overall layout of the entire machine, and the component element. When different elements share a node, the node is numbered according to the element that was built first. If different elements already have nodes at the same location, the larger number will be retained by default when the software is used to share the node.

[0065] The unit number is 8 digits long, and the sixth digit is used to describe the unit type.

[0066] The material number is 8 digits, the 3rd digit is subordinate to the 2nd digit, and the last 2 digits are reserved for future use;

[0067] The attribute number is 8 bits long. When multiple units share an attribute, the attribute number is the minimum value among the unit numbers.

[0068] In the present invention, the scale of the model is reduced, and the model adopts the simplified principle of one-dimensional / two-dimensional unit composite to improve the calculation efficiency without changing the basic geometric characteristics, force transmission path, stiffness and mass characteristics of the structure.

[0069] The present invention will be described in more detail below through examples.

[0070] Example 1:

[0071] like Figure 2 As shown, this embodiment proposes a finite element modeling method for a composite material ultra-large aspect ratio UAV, including:

[0072] (1) Divide the structural components into sub-models according to the overall layout of the whole aircraft to ensure that the sub-model interface is clear;

[0073] (2) Determine the type of components in the component, determine the model simplification principle and establish units;

[0074] (3) Create materials; determine the material type; in some cases, materials need to be simplified;

[0075] (4) Create and assign unit attributes; assign the built materials to the corresponding units according to the area;

[0076] (5) Refer to the numbering rules and renumber the nodes / elements / materials / attributes in the model;

[0077] (6) Debug the sub-model until it runs smoothly and create the finite element model of each structural component;

[0078] (7) Integrate all sub-models to obtain the full-machine finite element model, and debug it until it runs smoothly;

[0079] The specific implementation steps of this method include:

[0080] (1) Divide the structural components into sub-models according to the overall layout of the whole machine to ensure that the sub-model interface is clear.

[0081] Based on the overall layout definition of the composite material ultra-large aspect ratio UAV, the finite element model of the entire aircraft is divided into multiple sub-models based on components such as the wing section, fuselage, tail, and landing gear. The components in all sub-models follow the same simplification principle to achieve standardized and non-differential modeling between sub-models.

[0082] (2) Determine the type of components in the component, select the model simplification principle and establish the unit;

[0083] a) If the element structure is simple and can be loaded by tension or compression, it can be simplified to shell element;

[0084] For example, the wing main beam; as the most important load-bearing structure of a large aspect ratio UAV, the rationality of its simplification directly affects the overall stiffness accuracy of the model; taking a thin-walled circular tube long beam made of composite material as an example, a two-dimensional quadrilateral shell element is used for finite element modeling to simplify the process; the number of circumferential elements mainly controls the element size, that is, first determine the equally divided arc value to obtain the number of circumferential nodes, calculate the arc length corresponding to the arc value, control the slenderness ratio to 0.5~1.5 to determine the number of spanwise nodes, and divide the regular quadrilateral shell elements; such as Figure 2 , a unit is divided every 6° in the circumferential direction, with a total of 60 nodes, and the spanwise grid is divided with a slenderness ratio of 1;

[0085] b) If the element structure is irregular and the model has multi-dimensional combinations of surfaces and lines after geometric pre-processing, a composite simplification using shell elements and beam elements is used;

[0086] For example, the wing ribs; the wing ribs serve as the supporting structure of the airfoil shape, transferring aerodynamic loads to the main beam, and a simplified composite model of shell elements and beam elements is used; the wing rib webs use triangular / quadrilateral shell elements, and the element size is determined according to the main beam grid; the wing rib flanges use beam elements, and the size is the side length of the wing rib web element; the wing rib braces use beam elements, and the beam element nodes are the structural nodes;

[0087] Such as connecting flanges and fasteners; connecting flanges and fasteners; using shell elements and beam elements to simplify the model; flanges use triangular / quadrilateral shell elements, and the element size is determined according to the specific three-dimensional structure; fasteners use beam elements, and the beam element nodes are the nodes formed at the hole center after the connection hole is filled;

[0088] c) If the element structure is simple and can only bear tension, it can be simplified into a membrane element;

[0089] For example, the wing skin; considering the stress characteristics of the skin structure, which is tension but not compression, a two-dimensional membrane element is used to simplify the model, and the element size is determined according to the rib grid;

[0090] d) The connection joints of non-primary load-bearing components are simplified to RBE2 rigid constraint units;

[0091] In the full aircraft finite element model, the connectors / joints of non-primary load-bearing structures are simplified to RBE2 rigid constraint elements. The strength of such structures requires a separate detailed model verification.

[0092] (3) Create materials; determine the material type; in some cases, the materials need to be simplified;

[0093] a) Isotropic materials and orthotropic materials;

[0094] Performance parameters are determined based on performance test data;

[0095] b) laminated composite materials;

[0096] Finite element software is used to lay out the layers based on the actual layup information. The finite element software automatically calculates the comprehensive performance of the laminated composite material. When the actual performance of the laminate structure is measured, the performance of the single layer material should be appropriately corrected.

[0097] c) simplifying various layup methods into a unified laminated composite material;

[0098] In a small area, if there are variations in the layup of a composite structural component, taking into account both computational efficiency and accuracy, the area with varying layups is simplified. The simplified principle for unifying the laminated composite materials with different layups within the area is: based on the proportion of all layup methods within the area, the laminated composite material in that area is created using the layup method that covers the largest structural area.

[0099] The layup method of the laminated composite thin-walled circular tube beam is shown in Table 1. The layup changes along the circumferential direction and span direction of the circular tube. Figure 3 The attribute partitioning in the unification of the material layer in each area, such as the simplified material layer of attribute zone H3Z2 is shown in Table 2;

[0100] Table 1 Examples of laminated composite materials for thin-walled circular tube beams

[0101]

[0102] Table 2 Simplified example of ply layup in the H3Z2 property consistent zone of thin-walled circular tube beam

[0103] Number of layers Single layer material Laying angle / ° Initial number of layers 1 a 90 1 2 a 90 2 3 b 90 4 4 b 40 5 5 b 90 6 6 b 49 7 7 b 45 8 8 b 90 12 9 b 90 14 10 a 90 16 11 a 90 17

[0104] d) equivalent isotropic material of laminated composites;

[0105] If the actual structure simplified as beam elements in the model is a laminated composite material, use the comprehensive performance calculated by finite element software to create an isotropic material with the equivalent elastic modulus of the beam element in the axial force direction;

[0106] (4) Create and assign unit attributes; assign the built materials to the corresponding units according to the area;

[0107] a) Beam element: define the cross-sectional size information of the beam element according to the actual cross-sectional shape, and assign the cross-sectional information and material information to the beam element;

[0108] b) Shell / membrane element: assign material information to the corresponding shell / membrane element according to the actual structure;

[0109] c) simplifying various layup methods into a unified unit of laminated composite materials;

[0110] In a small range, if there are changes in the layup of composite structural components, taking into account both computational efficiency and accuracy, the layup change area is simplified, and the laminated composite materials with different layups are equivalent to the same material. The regional division principle is as follows: based on the actual layup situation, in the direction where the layup change has a relatively obvious impact on the structural performance, Nc = 5 to 8 attribute areas are divided; in the direction perpendicular to it, Nk = 1 / 10 to 1 / 8 times the aspect ratio of the attribute area are divided. That is, within this range, the model is divided into Nc*Nk attribute areas in total; the number of attribute partitions can be increased or decreased when necessary; the elements in the same attribute area are assigned the material information created after simplification;

[0111] like Figure 2 In the circumferential region H, every three rows of elements are grouped together. Since the laminated composite material plies in this example are symmetrical about the cross-section center, the circumferential region is divided into five groups of attribute-consistent regions. In the spanwise region Z, every 600 mm is grouped together, and the spanwise region is divided into three groups of attribute-consistent regions. That is, within this range, the model is divided into 15 attribute regions.

[0112] (5) Refer to the numbering rules and renumber the nodes / elements / materials / attributes in the model;

[0113] The principle of numbering the whole aircraft finite element model is to avoid duplication of node / element / material numbers between different sub-models. It is recommended to use the sub-model as the first digit of the number. The specific numbering rules are described as follows:

[0114] a) Submodel; the submodel file name should reflect the submodel number, design round number, modeling time, and other information; the calculation file submitted for the submodel should have the same name, and the processing condition abbreviation can be used to distinguish different calculation load conditions when necessary; for example, P_5_01_20240624 is the submodel file for the first design round of the wing component of the P project, created on June 24, 2024;

[0115] b) Node numbers are 8 digits long. The first three digits represent the component sub-model, the component's location in the overall aircraft layout, and the component element, respectively, which can be used to roughly locate the node. The fourth and fifth digits follow a certain pattern, such as wings can be sorted by ribs, and fuselages can be sorted by frame positions. Recommended rules and examples are shown in Table 3:

[0116] Table 3 Node numbering rules and examples

[0117]

[0118]

[0119] When different elements share a node, the node is numbered according to the element that was created first. If different elements already have nodes at the same location, the larger number will be retained by default when the software is used to share the node.

[0120] c) The unit number is 8 digits. The numbering rules are similar to the node numbering rules. The sixth digit is changed to indicate the unit type. The recommended rules and examples are shown in Table 4:

[0121] Table 4 Unit numbering rules and examples

[0122]

[0123] d) The material number is 8 digits; the 3rd digit is subordinate to the 2nd digit, and the last two digits are reserved for future use; the recommended rules and examples are shown in Table 5:

[0124] Table 5 Material numbering rules and examples

[0125]

[0126] e) The attribute number is 8 bits long and is the same as the unit number. When multiple units share the same attribute, the attribute number is the minimum value among the unit numbers. An example is shown in Table 6:

[0127] Table 6 Example of attribute numbers

[0128]

[0129] f) The local coordinate system number is 3 digits; the recommended rules and examples are shown in Table 7:

[0130] Table 7 Local coordinate system numbering rules and examples

[0131]

[0132] g) Flexible use of grouping functions: During the modeling process, use the grouping function to divide the model into multiple groups according to needs, which is conducive to improving modeling efficiency. For example, group by structural elements, group by groups with the same attributes, group by key areas when extracting calculation results, etc. The naming rules of the groups can refer to the numbering rules mentioned above and should be intuitive and concise;

[0133] (6) Debug the sub-model until it runs smoothly and create the finite element model of each structural component;

[0134] (7) Integrate all sub-models to obtain the finite element model of the entire aircraft, and debug it until it runs smoothly; at this point, the modeling work of the finite element model of the entire composite material ultra-large aspect ratio UAV is completed.

[0135] Example 2:

[0136] This embodiment provides a finite element modeling method for a composite material ultra-large aspect ratio UAV. Figure 1 As shown, including:

[0137] According to the overall layout of the whole aircraft, the UAV is divided into multiple structural component sub-models;

[0138] Determine the structural type of the components in the UAV structure, select the sub-model simplification principle and establish the unit;

[0139] Create materials, create and assign unit properties based on the created materials;

[0140] Renumber each sub-model and its nodes / elements / materials / attributes according to the numbering principle;

[0141] Debug each sub-model until it runs smoothly;

[0142] Integrate all sub-models to obtain the full aircraft finite element model, and debug it until it runs smoothly;

[0143] In this embodiment, multiple structural component sub-models are obtained by dividing the UAV into components;

[0144] UAVs are divided into wing sections, fuselage, tail wing and landing gear according to their components;

[0145] The structural components in all sub-models follow the same simplification principle to achieve standardized and indiscriminate modeling among sub-models;

[0146] In this embodiment, determining the structural type of the components of the drone structural parts, selecting the sub-model simplification principle and establishing the unit includes:

[0147] If the structural form of the component elements is simple and can be loaded by tension or compression, it can be simplified to shell elements;

[0148] If the structural form of the constituent elements is irregular and the model has multi-dimensional combinations of surfaces and lines after geometric pre-processing, a composite simplification using shell elements and beam elements is used;

[0149] If the structural form of the component is simple and can only bear tension, it can be simplified into a membrane unit;

[0150] If the component is a connection joint of a non-primary load-bearing component, it is simplified to an RBE2 rigid constraint element;

[0151] In this embodiment, creating materials includes:

[0152] Determine the material type, determine whether to simplify the material based on the material type, and simplify the material that needs to be simplified;

[0153] Determine whether to simplify the material based on the material type. The simplification of materials that need to be simplified includes:

[0154] If it is an isotropic material or an orthotropic material, no simplification is done and its performance parameters are determined based on the performance test data;

[0155] If it is a laminated composite material, it is simplified to an isotropic material. The simplification principle is as follows: the comprehensive performance of the laminated composite material is calculated using finite element software, and the isotropic material is created using the equivalent elastic modulus of the beam element in the direction of the axial force;

[0156] If there are multiple layup methods, simplify them into a unified laminated composite material. The simplification principle is: based on the proportion of all layup methods in the region, use the layup method that covers the largest structural area to create a regional laminated composite material;

[0157] In this embodiment, creating and assigning unit attributes based on the created material includes:

[0158] Define the cross-sectional size information of the beam element according to the actual cross-sectional shape, and assign the cross-sectional information and material information to the beam element;

[0159] Assign material information to the corresponding shell / membrane elements according to the actual structure;

[0160] According to the actual ply situation, in the direction where the ply change has a relatively obvious impact on the structural performance, Nc = 5 to 8 attribute zones are divided. In the direction perpendicular to it, Nk = 1 / 10 to 1 / 8 times the aspect ratio attribute zones are divided. That is, within this range, the model is divided into Nc*Nk attribute zones in total.

[0161] In this embodiment, the numbering principle is: the sub-model is used as the first identifier of the number to avoid duplication of node / element / material numbers between different sub-models;

[0162] In this embodiment, renumbering each sub-model and its nodes / elements / materials / attributes includes:

[0163] The file name of the sub-model reflects the sub-model number, design round number and modeling time;

[0164] The node number is 8 digits. The first three digits represent the structural component submodel, the structural component's location in the overall layout of the entire machine, and the component element. When different elements share a node, the node is numbered according to the element that was built first. If different elements already have nodes at the same location, the larger number will be retained by default when the software is used to share the node.

[0165] The unit number is 8 digits long, and the sixth digit is used to describe the unit type.

[0166] The material number is 8 digits, the 3rd digit is subordinate to the 2nd digit, and the last 2 digits are reserved for future use;

[0167] The attribute number is 8 bits long. When multiple units share an attribute, the attribute number is the minimum value among the unit numbers.

[0168] Example 3:

[0169] This embodiment provides a finite element modeling device for a composite material ultra-high aspect ratio UAV, comprising:

[0170] The sub-model division module is used to divide the UAV into multiple structural component sub-models according to the overall layout of the whole aircraft;

[0171] The unit establishment module is used to determine the structural type of the components in the UAV structure, select the sub-model simplification principle and establish the unit;

[0172] Material creation module, used to create materials, create and assign unit properties based on the created materials;

[0173] The renumbering module is used to renumber each sub-model and its nodes / elements / materials / attributes according to the numbering principle;

[0174] Sub-model debugging module, used to debug each sub-model until it runs smoothly;

[0175] The finite element model debugging module is used to integrate all sub-models to obtain the full aircraft finite element model and debug it until it runs smoothly;

[0176] According to the overall layout of the whole aircraft, the UAV is divided into multiple structural component sub-models;

[0177] Determine the structural type of the components in the UAV structure, select the sub-model simplification principle and establish the unit;

[0178] Create materials, create and assign unit properties based on the created materials;

[0179] Renumber each sub-model and its nodes / elements / materials / attributes according to the numbering principle;

[0180] Debug each sub-model until it runs smoothly;

[0181] Integrate all sub-models to obtain the full aircraft finite element model, and debug it until it runs smoothly;

[0182] In this embodiment, multiple structural component sub-models are obtained by dividing the UAV into components;

[0183] UAVs are divided into wing sections, fuselage, tail and landing gear according to their components;

[0184] The structural components in all sub-models follow the same simplification principle to achieve standardized and indiscriminate modeling among sub-models;

[0185] In this embodiment, determining the structural type of the components of the drone structural parts, selecting the sub-model simplification principle and establishing the unit includes:

[0186] If the structural form of the component elements is simple and can be loaded by tension or compression, it can be simplified to shell elements;

[0187] If the structural form of the constituent elements is irregular and the model has multi-dimensional combinations of surfaces and lines after geometric pre-processing, a composite simplification using shell elements and beam elements is used;

[0188] If the structural form of the component is simple and can only bear tension, it can be simplified into a membrane unit;

[0189] If the component is a connection joint of a non-primary load-bearing component, it is simplified to an RBE2 rigid constraint element;

[0190] In this embodiment, creating materials includes:

[0191] Determine the material type, determine whether to simplify the material based on the material type, and simplify the material that needs to be simplified;

[0192] Determine whether to simplify the material based on the material type. The simplification of materials that need to be simplified includes:

[0193] If it is an isotropic material or an orthotropic material, no simplification is done and its performance parameters are determined based on the performance test data;

[0194] If it is a laminated composite material, it is simplified to an isotropic material. The simplification principle is as follows: the comprehensive performance of the laminated composite material is calculated using finite element software, and the isotropic material is created using the equivalent elastic modulus of the beam element in the direction of the axial force;

[0195] If there are multiple layup methods, simplify them into a unified laminated composite material. The simplification principle is: based on the proportion of all layup methods in the region, use the layup method that covers the largest structural area to create a regional laminated composite material;

[0196] In this embodiment, creating and assigning unit attributes based on the created material includes:

[0197] Define the cross-sectional size information of the beam element according to the actual cross-sectional shape, and assign the cross-sectional information and material information to the beam element;

[0198] Assign material information to the corresponding shell / membrane elements according to the actual structure;

[0199] According to the actual ply situation, in the direction where the ply change has a relatively obvious impact on the structural performance, Nc = 5 to 8 attribute zones are divided. In the direction perpendicular to it, Nk = 1 / 10 to 1 / 8 times the aspect ratio attribute zones are divided. That is, within this range, the model is divided into Nc*Nk attribute zones in total.

[0200] In this embodiment, the numbering principle is: the sub-model is used as the first identifier of the number to avoid duplication of node / element / material numbers between different sub-models;

[0201] In this embodiment, renumbering each sub-model and its nodes / elements / materials / attributes includes:

[0202] The file name of the sub-model reflects the sub-model number, design round number and modeling time;

[0203] The node number is 8 digits. The first three digits represent the structural component submodel, the structural component's location in the overall layout of the entire machine, and the component element. When different elements share a node, the node is numbered according to the element that was built first. If different elements already have nodes at the same location, the larger number will be retained by default when the software is used to share the node.

[0204] The unit number is 8 digits long, and the sixth digit is used to describe the unit type.

[0205] The material number is 8 digits, the 3rd digit is subordinate to the 2nd digit, and the last 2 digits are reserved for future use;

[0206] The attribute number is 8 bits long. When multiple units share an attribute, the attribute number is the minimum value among the unit numbers.

[0207] Example 4:

[0208] An embodiment of the present invention provides an electronic device including a memory and a processor, including:

[0209] a memory storing executable instructions;

[0210] The processor runs the executable instructions in the memory to implement a finite element modeling method for a composite material ultra-large aspect ratio UAV.

[0211] The memory is used to store non-transitory computer-readable instructions. Specifically, the memory may include one or more computer program products, which may include various forms of computer-readable storage media, such as volatile memory and / or non-volatile memory. The volatile memory may include, for example, random access memory (RAM) and / or cache memory. The non-volatile memory may include, for example, read-only memory (ROM), a hard disk, flash memory, etc.

[0212] The processor may be a central processing unit (CPU) or other forms of processing units having data processing capabilities and / or instruction execution capabilities, and may control other components in the electronic device to perform desired functions. In one embodiment of the present invention, the processor is used to run the computer-readable instructions stored in the memory.

[0213] Those skilled in the art should understand that in order to solve the technical problem of how to obtain a good user experience, this embodiment may also include well-known structures such as a communication bus and an interface, and these well-known structures should also be included in the scope of protection of the present invention.

[0214] For detailed description of this embodiment, please refer to the corresponding description in the aforementioned embodiments, which will not be repeated here.

[0215] Embodiment 5:

[0216] An embodiment of the present invention provides a computer-readable storage medium storing a computer program. When the computer program is executed by a processor, a finite element modeling method for a composite material ultra-large aspect ratio unmanned aerial vehicle is implemented.

[0217] The computer-readable storage medium according to an embodiment of the present invention stores non-transitory computer-readable instructions, which, when executed by a processor, execute all or part of the steps of the aforementioned methods of the various embodiments of the present invention.

[0218] The above-mentioned computer-readable storage media include, but are not limited to, optical storage media (e.g., CD-ROMs and DVDs), magneto-optical storage media (e.g., MOs), magnetic storage media (e.g., magnetic tapes or mobile hard disks), media with built-in rewritable non-volatile memory (e.g., memory cards), and media with built-in ROM (e.g., ROM cartridges).

[0219] The finite element modeling method for a composite material ultra-large aspect ratio UAV proposed in an embodiment of the present invention provides corresponding simplification principles for different structural components in the aircraft. While ensuring that the basic geometric characteristics, force transmission paths, stiffness and mass characteristics of the structure are not changed after simplification, it unifies the modeling standards, regulates the element numbering, reduces the model size, increases the readability of the calculation file and improves the calculation efficiency.

[0220] While various embodiments of the present invention have been described above, the above description is intended to be illustrative, not exhaustive, and not limited to the disclosed embodiments. Many modifications and variations will be apparent to those skilled in the art without departing from the scope and spirit of the described embodiments.

Claims

1. A finite element modeling method for a composite material ultra-large aspect ratio UAV, characterized in that: include: According to the overall layout of the whole aircraft, the UAV is divided into multiple structural component sub-models; Determining the structural type of the components of the UAV structural components, selecting the sub-model simplification principle and establishing units; Create materials, create and assign unit properties based on the created materials; Renumber each sub-model and its nodes / elements / materials / attributes according to the numbering principle; Debug each sub-model until it runs smoothly; Integrate all sub-models to obtain the full aircraft finite element model, and debug it until it runs smoothly; Determining the structural type of the components of the UAV structural components, selecting the sub-model simplification principle and establishing the unit includes: If the structural form of the component elements is simple and can be loaded by tension or compression, it can be simplified to shell elements; If the structural form of the constituent elements is irregular and the model has multi-dimensional combinations of surfaces and lines after geometric pre-processing, a composite simplification using shell elements and beam elements is used; If the structural form of the component is simple and can only bear tension, it can be simplified into a membrane unit; If the component is a connection joint of a non-primary load-bearing component, it is simplified to an RBE2 rigid constraint element; Create and assign unit properties based on the created material, including: Define the cross-sectional size information of the beam element according to the actual cross-sectional shape, and assign the cross-sectional information and material information to the beam element; Assign material information to the corresponding shell / membrane elements according to the actual structure; According to the actual ply situation, in the direction where the ply change has a relatively obvious impact on the structural performance, Nc = 5 to 8 attribute zones are divided. In the direction perpendicular to it, Nk = 1 / 10 to 1 / 8 times the aspect ratio attribute zones are divided. That is, within this range, the model is divided into Nc*Nk attribute zones in total. The numbering principle is: use the sub-model as the first digit of the number to avoid duplication of node / element / material numbers between different sub-models; Renumbering of each sub-model and its nodes / elements / materials / properties includes: The file name of the sub-model reflects the sub-model number, design round number and modeling time; The node number is 8 digits. The first three digits represent the structural component submodel, the structural component's location in the overall layout of the entire machine, and the component element. When different elements share a node, the node is numbered according to the element that was built first. If different elements already have nodes at the same location, the larger number will be retained by default when the software is used to share the node. The unit number is 8 digits long, and the sixth digit is used to describe the unit type. The material number is 8 digits, the 3rd digit is subordinate to the 2nd digit, and the last 2 digits are reserved for future use; The attribute number is 8 bits long. When multiple units share an attribute, the attribute number is the minimum value among the unit numbers.

2. The method according to claim 1, characterized in that The multiple structural component sub-models are obtained by dividing the UAV into components; The UAV is divided into wing sections, fuselage, tail wing and landing gear according to its components; The structural components in all sub-models follow the same simplification principle to achieve standardized and non-differentiated modeling among the sub-models.

3. The method according to claim 2, characterized in that Creation materials include: Determine the material type, determine whether to simplify the material based on the material type, and simplify the material that needs to be simplified; Determine whether to simplify the material based on the material type. The simplification of materials that need to be simplified includes: If it is an isotropic material or an orthotropic material, no simplification is done and its performance parameters are determined based on the performance test data; If it is a laminated composite material, it is simplified to an isotropic material. The simplification principle is as follows: the comprehensive performance of the laminated composite material is calculated using finite element software, and the isotropic material is created using the equivalent elastic modulus of the beam element in the direction of the axial force; If there are multiple layup methods, they can be simplified into a unified laminated composite material. The simplification principle is: based on the proportion of all layup methods in the region, the layup method with the largest coverage area is used to create a regional laminated composite material.

4. A finite element modeling device for composite material ultra-large aspect ratio UAV, characterized in that: include: The sub-model division module is used to divide the UAV into multiple structural component sub-models according to the overall layout of the whole aircraft; A unit establishment module, configured to determine the structural type of the components of the UAV structural components, select the sub-model simplification principle and establish a unit; Material creation module, used to create materials, create and assign unit properties based on the created materials; The renumbering module is used to renumber each sub-model and its nodes / elements / materials / attributes according to the numbering principle; Sub-model debugging module, used to debug each sub-model until it runs smoothly; The finite element model debugging module is used to integrate all sub-models to obtain the full aircraft finite element model and debug it until it runs smoothly; Determining the structural type of the components of the UAV structural components, selecting the sub-model simplification principle and establishing the unit includes: If the structural form of the component elements is simple and can be loaded by tension or compression, it can be simplified to shell elements; If the structural form of the constituent elements is irregular and the model has multi-dimensional combinations of surfaces and lines after geometric pre-processing, a composite simplification using shell elements and beam elements is used; If the structural form of the component is simple and can only bear tension, it can be simplified into a membrane unit; If the component is a connection joint of a non-primary load-bearing component, it is simplified to an RBE2 rigid constraint element; Create and assign unit properties based on the created material, including: Define the cross-sectional size information of the beam element according to the actual cross-sectional shape, and assign the cross-sectional information and material information to the beam element; Assign material information to the corresponding shell / membrane elements according to the actual structure; According to the actual ply situation, in the direction where the ply change has a relatively obvious impact on the structural performance, Nc = 5 to 8 attribute zones are divided. In the direction perpendicular to it, Nk = 1 / 10 to 1 / 8 times the aspect ratio attribute zones are divided. That is, within this range, the model is divided into Nc*Nk attribute zones in total. The numbering principle is: use the sub-model as the first digit of the number to avoid duplication of node / element / material numbers between different sub-models; Renumbering of each sub-model and its nodes / elements / materials / properties includes: The file name of the sub-model reflects the sub-model number, design round number and modeling time; The node number is 8 digits. The first three digits represent the structural component submodel, the structural component's location in the overall layout of the entire machine, and the component element. When different elements share a node, the node is numbered according to the element that was built first. If different elements already have nodes at the same location, the larger number will be retained by default when the software is used to share the node. The unit number is 8 digits long, and the sixth digit is used to describe the unit type. The material number is 8 digits, the 3rd digit is subordinate to the 2nd digit, and the last 2 digits are reserved for future use; The attribute number is 8 bits long. When multiple units share an attribute, the attribute number is the minimum value among the unit numbers.

5. An electronic device, characterized in that: The electronic device comprises: a memory storing executable instructions; A processor, wherein the processor runs the executable instructions in the memory to implement the finite element modeling method for a composite material ultra-large aspect ratio UAV according to any one of claims 1 to 3.

6. A computer-readable storage medium, characterized in that The computer-readable storage medium stores a computer program, which, when executed by a processor, implements the finite element modeling method for a composite material ultra-large aspect ratio UAV according to any one of claims 1 to 3.

Images