Finite element modeling method and device for composite unmanned aerial vehicle with ultra-high aspect ratio

Abstract

The invention relates to the technical field of unmanned aerial vehicle design, and particularly discloses a finite element modeling method and device for a composite unmanned aerial vehicle with an ultra-high aspect ratio, and the method comprises the steps: dividing the unmanned aerial vehicle into a plurality of structural component sub-models according to the overall layout of the whole unmanned aerial vehicle; judging the structure type of the component elements of the unmanned aerial vehicle structure component, selecting a sub-model simplification principle and establishing a unit; creating a material, and creating and endowing unit attributes based on the created material; according to a numbering principle, each sub-model and the node / unit / material / attribute thereof are numbered again; debugging each sub-model until the sub-model runs smoothly; and integrating all the sub-models to obtain a full-aircraft finite element model, and debugging the full-aircraft finite element model until the full-aircraft finite element model runs smoothly. According to the method, corresponding simplification principles are provided for different structural components of an aircraft, and on the premise of ensuring that basic geometric characteristics, force transmission routes, rigidity, quality characteristics and the like of the structure are not changed after simplification, modeling standards and rule element numbers are unified, the model scale is reduced, and the calculation efficiency is improved.

Description

Technical Field

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

[0002] Composite material super-large aspect ratio UAV is lightweight and achieves design weight closure by making full use of the designability of laminated composite materials. It is mainly affected by gust loads. When establishing a full-machine finite element model for it, in addition to the general modeling principles, it is necessary to establish a full-machine finite element model of the static strength of the composite material super-large aspect ratio UAV based on its unique structural design characteristics. This model must be suitable for the analysis and calculation of the structural strength, stiffness and stability of the UAV, 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 reduction of model scale, which leads to poor readability of model parameters, weak operability, and low calculation efficiency, greatly affecting 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 view of the shortcomings of the prior art, the present invention provides a finite element modeling method and device for a composite material ultra-large aspect ratio UAV. The method provides corresponding simplification principles for different structural components in the aircraft. On 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 regular, the model scale is reduced, the readability of the calculation file is increased, and the calculation efficiency is improved.

[0006] In order to achieve the above object, a first aspect of the present invention provides a finite element modeling method for a composite material ultra-large 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] Determine the structural type of the components in the UAV structural component, select the sub-model simplification principle and establish a unit;

[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, used to determine the structural type of the components in the UAV structural components, select the sub-model simplification principle and establish a unit;

[0016] The material creation module is used to create materials and to create and assign unit attributes based on the created materials;

[0017] The renumbering module is used to renumber each sub-model and its nodes / units / 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 whole machine, and debug it until it runs smoothly.

[0020] A third aspect of the present invention provides an electronic device, the 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 described in 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 a composite material ultra-large aspect ratio UAV as 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 the present invention provides corresponding simplification principles for different structural components in the aircraft. On 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 regular, the model scale is reduced, the readability of the calculation file 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 non-differentiated modeling for multiple people; the 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.

[0027] (3) The finite element modeling method for the composite material ultra-large aspect ratio UAV proposed in the present 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 schematic 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 The present invention is a flowchart 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 As 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 the 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, the modeling standards are unified, the element numbers are regular, the model scale is reduced, the readability of the calculation file is increased, and the calculation efficiency is improved.

[0042] According to the present invention, a plurality of structural component sub-models are obtained by dividing the drone by 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, the structural type of the components in the UAV structural parts is determined, the sub-model simplification principle is selected and the unit is established, including:

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

[0047] If the structural form of the constituent elements is irregular, and after geometric pre-processing, the model has a multi-dimensional combination of surfaces and lines, the shell element and beam element composite simplification shall be adopted;

[0048] If the structural form of the component elements 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 is 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, thus improving modeling efficiency and achieving non-differentiated modeling among multiple people; regular element numbering is adopted, and the numbering of nodes / units / materials / attributes of the input files submitted for calculation is regular and orderly, with good readability and strong operability.

[0051] According to the present invention, creating materials 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. Simplify the material that needs to be simplified, including:

[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 laid out 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 laying methods, they can be simplified into a unified laminated composite material. The simplification principle is: according to the proportion of all laying methods in the region, the laying method with the largest covering structure area is adopted 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 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=5~8 attribute zones are divided, and in the direction perpendicular to it, Nk=1 / 10~1 / 8 times the length-to-width 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 digit of the numbering to avoid duplication of node / unit / material numbers 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 sub-model of the structural component, the location of the structural component in the overall layout of the whole 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 position, the larger number value is 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 long, with the 3rd digit being subordinate to the 2nd digit, and the last two digits being 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 combination 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 by way of examples.

[0070] Embodiment 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 principles 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 finite element model of the entire aircraft, 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, and ensure that the sub-model interface is clear.

[0081] According to the overall layout definition of composite material ultra-large aspect ratio UAV, the finite element model of the whole aircraft is divided into multiple sub-models according to the wing section, fuselage, tail, landing gear and other components. The components in all sub-models follow the same simplification principle to achieve standardized and indiscriminate 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 of a single type and can be loaded in tension or compression, it can be simplified to a shell element;

[0084] Such as the wing main beam; the main beam is the most important load-bearing structure of the ultra-large aspect ratio UAV, and the rationality of its simplification directly affects the overall stiffness accuracy of the model; taking the composite thin-walled circular tube long beam as an example, the finite element modeling adopts two-dimensional quadrilateral shell element simplification; the number of circumferential elements mainly controls the unit size, that is, first determine the number of circumferential nodes obtained by equally dividing the arc value, calculate the arc length corresponding to the arc value, control the slenderness ratio of 0.5 to 1.5 to determine the number of spanwise nodes, and divide the regular quadrilateral shell elements; such as Figure 2 , one 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, after geometric pre-processing, the model has multi-dimensional combinations of surfaces and lines, and shell elements and beam elements are used for composite simplification;

[0086] Such as the wing ribs; the wing ribs are the supporting structure of the airfoil shape, transferring the aerodynamic load to the main beam, and a composite simplified model of shell elements and beam elements is adopted; the wing rib web adopts triangular / quadrilateral shell elements, and the element size is determined according to the main beam grid; the wing rib flange adopts beam elements, and the size is the side length of the wing rib web element; the wing rib brace adopts 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; the flange uses triangular / quadrilateral shell elements, and the element size is determined according to the specific three-dimensional structure; the fasteners use beam elements, and the beam element nodes are the nodes formed at the center of the hole 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 unit;

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

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

[0091] In the finite element model of the whole aircraft, the connectors / joints of the non-main load-bearing structures are simplified into RBE2 rigid constraint units. The strength of such structures needs to be verified by separately establishing detailed models;

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

[0093] a) Isotropic materials and orthotropic materials;

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

[0095] b) laminated composite materials;

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

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

[0098] In a small area, if there are changes in the layup of composite structural components, taking into account both computational efficiency and accuracy, simplify the layup change area, and unify the laminated composite materials with different layups in the area. The simplification principle is: according to the proportion of all layup methods in the area, the layup method with the largest coverage of the structural area is adopted to create the laminated composite material in the 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 unifies the material layup in each area. For example, the simplified result of the material layup in the 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 Example of simplified plywork 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, the comprehensive performance calculated by finite element software is used 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 dimension 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 unit: assign material information to corresponding shell / membrane unit 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 ply of composite structural components, taking into account the calculation efficiency and accuracy, simplify the ply change area, and the laminated composite materials with different plies are equivalent to the same material. The principle of regional division is as follows: according to the actual ply situation, in the direction where the ply change affects the structural performance relatively obviously, divide Nc = 5 to 8 attribute areas; in the direction perpendicular to it, divide Nk = 1 / 10 to 1 / 8 times the aspect ratio attribute area, that is, within this range, the model is divided into Nc*Nk attribute areas in total; if necessary, the number of attribute partitions can be increased or decreased; for the units in the same attribute area, the material information created after simplification is given;

[0111] like Figure 2 , every three rows of units in the annular region H are taken as a group. Since the laminated composite material plies in the example are symmetrical about the cross-section center, the annular region is divided into 5 groups of attribute consistent areas; every 600mm in the spanwise region Z is taken as a group, and the spanwise region is divided into 3 groups of attribute consistent areas. That is, within this range, the model is divided into 15 attribute areas;

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

[0113] The numbering principle of the whole aircraft finite element model is to avoid duplication of node / unit / material numbers between different sub-models. It is recommended to use the sub-model as the first digit of the numbering. 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 by the submodel shall have the same name, and the processing condition abbreviation can be used to distinguish different calculation load conditions when necessary; such as P_5_01_20240624, which is the submodel file of the first round of design of the wing component of the P project, created on June 24, 2024;

[0115] b) The node number is 8 digits; the first three digits are the component sub-model, the component is located in the overall layout of the whole aircraft, and the component element, which can roughly locate the node; the 4th and 5th digits follow a certain rule, such as the wing can be sorted in sequence by wing ribs, and the fuselage can be sorted in sequence by frame station position, etc. The 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 with the element that was built first; if different elements already have nodes at the same location, when the software is used to share the node, the larger number value is retained by default;

[0120] c) The unit number is 8 digits. The numbering rules are similar to the node numbering rules. The sixth digit of the number 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 one attribute, the attribute number is the minimum value in the unit numbers. An example is shown in Table 6:

[0127] Table 6 Example of attribute number

[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 function: In the process of modeling, the grouping function can be used to divide the model into multiple groups according to the needs, which is conducive to improving the modeling efficiency; for example, grouping by structural elements, grouping by the same attributes, grouping by the key areas when extracting calculation results, etc. The naming rules of the groups can refer to the above numbering rules 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; this completes the modeling of the finite element model of the entire composite material ultra-large aspect ratio UAV.

[0135] Embodiment 2:

[0136] This embodiment provides a finite element modeling method for a composite material ultra-large aspect ratio UAV, such as 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 finite element model of the whole aircraft, and debug it until it runs smoothly;

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

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

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

[0146] In this embodiment, the structural type of the components in the UAV structural parts is determined, the sub-model simplification principle is selected and the unit is established, including:

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

[0148] If the structural form of the constituent elements is irregular, and after geometric pre-processing, the model has a multi-dimensional combination of surfaces and lines, the shell element and beam element composite simplification shall be adopted;

[0149] If the structural form of the component elements 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 unit;

[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. Simplify the material that needs to be simplified, including:

[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 laid out 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: according to the proportion of all layup methods in the region, adopt the layup method with the largest coverage structure 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 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, and in the direction perpendicular to it, Nk = 1 / 10 to 1 / 8 times the length-to-width 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 digit of the numbering to avoid duplication of the numbering of nodes / units / materials between different sub-models;

[0162] In this embodiment, renumbering each sub-model and its nodes / units / 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 sub-model of the structural component, the location of the structural component in the overall layout of the whole 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 position, the larger number value is 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 long, with the 3rd digit being subordinate to the 2nd digit, and the last two digits being 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] Embodiment three:

[0169] This embodiment provides a finite element modeling device for a composite material ultra-large 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 structural parts, select the sub-model simplification principle and establish the unit;

[0172] The material creation module is used to create materials and to create and assign unit attributes based on the created materials;

[0173] The renumbering module is used to renumber each sub-model and its nodes / units / 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-machine 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 drone 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 indifferent modeling among sub-models;

[0185] In this embodiment, the structural type of the components in the drone structural parts is determined, the sub-model simplification principle is selected and the unit is established, including:

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

[0187] If the structural form of the constituent elements is irregular, and after geometric pre-processing, the model has a multi-dimensional combination of surfaces and lines, the shell element and beam element composite simplification shall be adopted;

[0188] If the structural form of the component elements 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 unit;

[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. Simplify the material that needs to be simplified, including:

[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 laid out 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: according to the proportion of all layup methods in the region, adopt the layup method with the largest coverage structure 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 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, and in the direction perpendicular to it, Nk = 1 / 10 to 1 / 8 times the length-to-width 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 digit of the numbering to avoid duplication of the numbering of nodes / units / materials between different sub-models;

[0201] In this embodiment, renumbering each sub-model and its nodes / units / 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 sub-model of the structural component, the location of the structural component in the overall layout of the whole 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 position, the larger number value is 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 long, with the 3rd digit being subordinate to the 2nd digit, and the last two digits being 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] Embodiment 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 the finite element modeling method of the composite material ultra-large aspect ratio UAV.

[0211] The memory is used to store non-temporary 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, for example, include random access memory (RAM) and / or cache memory (cache), etc. The non-volatile memory may, for example, include read-only memory (ROM), hard disk, flash memory, etc.

[0212] The processor may be a central processing unit (CPU) or other forms of processing units with 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 be able to understand that in order to solve the technical problem of how to obtain a good user experience, the present 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 protection scope of the present invention.

[0214] For detailed description of this embodiment, reference may be made to the corresponding descriptions in the aforementioned embodiments, which will not be repeated here.

[0215] Embodiment five:

[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 the embodiment of the present invention stores non-transitory computer-readable instructions, and when the non-transitory computer-readable instructions are executed by a processor, all or part of the steps of the above-mentioned methods of the embodiments of the present invention are executed.

[0218] The above-mentioned computer-readable storage media include, but are not limited to: optical storage media (e.g., CD-ROM and DVD), magneto-optical storage media (e.g., MO), magnetic storage media (e.g., magnetic tape or mobile hard disk), media with built-in rewritable non-volatile memory (e.g., memory card) and media with built-in ROM (e.g., ROM box).

[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. 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, the modeling standards are unified, the element numbering is regular, the model scale is reduced, the readability of the calculation file is increased, and the calculation efficiency is improved.

[0220] The embodiments of the present invention have been described above, and the above description is exemplary, not exhaustive, and is not limited to the disclosed embodiments. Many modifications and changes 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 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; Determine the structural type of the components in the UAV structural component, select the sub-model simplification principle and establish a unit; 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 finite element model of the entire aircraft, and debug it until it runs smoothly.

2. The method according to claim 1, characterized in that The multiple structural component sub-models are obtained by dividing the drone 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 Determining the structural type of the components in 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 in tension or compression, it can be simplified to shell elements; If the structural form of the constituent elements is irregular, and after geometric pre-processing, the model has a multi-dimensional combination of surfaces and lines, the shell element and beam element composite simplification shall be adopted; If the structural form of the component elements 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 unit.

4. The method according to claim 3, characterized in that Materials to create 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. Simplify the material that needs to be simplified, including: 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 laid out 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 laying methods, they can be simplified into a unified laminated composite material. The simplification principle is: according to the proportion of all laying methods in the region, the laying method with the largest covering structure area is adopted to create a regional laminated composite material.

5. The method according to claim 4, characterized in that Based on the created material, the following properties are created and assigned to the unit: 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 corresponding shell / membrane elements according to the actual structure; According to the actual ply situation, in the direction where the ply change affects the structural performance relatively obviously, Nc=5~8 attribute zones are divided, and in the direction perpendicular to it, Nk=1 / 10~1 / 8 times the length-to-width ratio attribute zones are divided. That is, within this range, the model is divided into Nc*Nk attribute zones in total.

6. The method according to claim 1, characterized in that The numbering principle is: use the sub-model as the first digit of the numbering to avoid duplication of node / unit / material numbers between different sub-models.

7. The method according to claim 6, characterized in that Renumbering of individual sub-models and their 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 sub-model of the structural component, the location of the structural component in the overall layout of the whole 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 position, the larger number value is 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 long, with the 3rd digit being subordinate to the 2nd digit, and the last two digits being 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.

8. 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, used to determine the structural type of the components in the UAV structural components, select the sub-model simplification principle and establish a unit; The material creation module is used to create materials and to create and assign unit attributes based on the created materials; The renumbering module is used to renumber each sub-model and its nodes / units / 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 finite element model of the whole machine, and debug it until it runs smoothly.

9. 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 7.

10. 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 7.

Images