Design method, device and storage medium for honeycomb sandwich structure of aircraft laminate

Abstract

The present invention discloses a design method, device and storage medium of an aircraft laminate honeycomb sandwich structure, comprising: obtaining body property data of an aircraft body structure with an initial laminate honeycomb sandwich structure, and constructing a finite element model based on the body property data; determining the actual load, actual displacement constraint and actual flight environment borne by the aircraft during service, applying the load, displacement constraint and flight environment to the finite element model based on the above conditions, and determining the stress state of the initial laminate honeycomb sandwich structure after applying the above conditions; determining multiple failure modes, and determining evaluation parameters of the initial laminate honeycomb sandwich structure in each failure mode based on the stress state; judging whether the initial laminate honeycomb sandwich structure meets the design requirements based on each evaluation parameter, and if so, determining the initial laminate honeycomb sandwich structure as a frozen laminate honeycomb sandwich structure, otherwise, performing structural optimization on the initial laminate honeycomb sandwich structure.

Description

Technical Field

[0001] The invention relates to the field of aircraft composite material design, and in particular to a design method, device and storage medium for an aircraft laminate honeycomb sandwich structure. Background Art

[0002] Composite laminates have the advantages of high strength, high stiffness, light weight, fatigue resistance, corrosion resistance, etc. Composite laminates / honeycomb sandwich structures (hereinafter referred to as "laminates / honeycomb sandwich structures") are a type of composite structure, which is widely used in the aircraft industry. Many commercial aircraft, helicopters and some military aircraft widely use laminates / honeycomb sandwich structures, such as wing fairings, control panels, landing gear doors, cockpit fairings, floor beams, wing leading and trailing edges, rudders, kitchens and rotor blades. However, due to the diversity of composite molding processes, it is particularly important to design laminates / honeycomb sandwich structures that meet the actual flight needs of aircraft.

[0003] At present, the aircraft laminate honeycomb sandwich structure designed based on human experience is usually subjected to static testing alone, and the aircraft laminate honeycomb sandwich structure is optimized based on the static test results. However, this optimization design method only considers the force applied by the external environment and cannot accurately simulate the actual flight conditions of the aircraft, resulting in low accuracy in the design of the aircraft laminate honeycomb sandwich structure. Summary of the invention

[0004] The present invention provides a design method, device and storage medium for an aircraft laminate honeycomb sandwich structure, which are mainly capable of improving the accuracy of the design of the aircraft laminate honeycomb sandwich structure.

[0005] According to a first aspect of the present invention, there is provided a design method for an aircraft laminate honeycomb sandwich structure, comprising:

[0006] Determine an initial laminate honeycomb sandwich structure of the laminate honeycomb sandwich to be designed, obtain fuselage property data of an aircraft fuselage structure with the initial laminate honeycomb sandwich structure, and construct a natural grid finite element model of the aircraft fuselage structure based on the fuselage property data, wherein the fuselage property data includes structural property data of the initial laminate honeycomb sandwich structure;

[0007] Determine the actual load, actual displacement constraint, and actual flight environment conditions borne by the actual aircraft during service, and determine a simulated load with the same load as the actual load, a simulated displacement constraint with the same constraint as the actual displacement constraint, and a simulated flight environment condition with the same environment as the actual flight environment condition;

[0008] Applying the simulated load, setting the simulated displacement constraint condition, and setting the simulated flight environment condition to the natural grid finite element model of the aircraft body structure, and determining the stress state data of the initial laminate honeycomb sandwich structure after applying the simulated load, setting the simulated displacement constraint condition, and setting the simulated flight environment condition;

[0009] Based on the stress state data, determining multiple failure modes of the initial laminate honeycomb sandwich structure, and based on the stress state data, determining structural design evaluation parameters of the initial laminate honeycomb sandwich structure under each of the failure modes;

[0010] Based on the structural design evaluation parameters and the structural property data, it is determined whether the initial laminate honeycomb sandwich structure meets the design requirements. If the initial laminate honeycomb sandwich structure meets the design requirements, the initial laminate honeycomb sandwich structure is determined as a frozen laminate honeycomb sandwich structure applied to an aircraft fuselage structure. Otherwise, the structural parameters of the initial laminate honeycomb sandwich structure are optimized.

[0011] According to a second aspect of the present invention, there is provided a design device for an aircraft laminate honeycomb sandwich structure, comprising:

[0012] An acquisition unit is used to determine an initial laminate honeycomb sandwich structure of the laminate honeycomb sandwich to be designed, and to acquire body property data of an aircraft body structure with the initial laminate honeycomb sandwich structure, and to construct a natural grid finite element model of the aircraft body structure based on the body property data, wherein the body property data includes structural property data of the initial laminate honeycomb sandwich structure;

[0013] The first determination unit is used to determine the actual load, actual displacement constraint condition, and actual flight environment condition borne by the actual aircraft during service, and determine a simulated load with the same load as the actual load, a simulated displacement constraint condition with the same constraint as the actual displacement constraint condition, and a simulated flight environment condition with the same environment as the actual flight environment condition;

[0014] an applying unit, used for applying the simulated load, setting the simulated displacement constraint condition, and setting the simulated flight environment condition to the natural grid finite element model of the aircraft body structure, and determining stress state data of the initial laminate honeycomb sandwich structure after applying the simulated load, setting the simulated displacement constraint condition, and setting the simulated flight environment condition;

[0015] A second determination unit is used to determine multiple failure modes of the initial laminate honeycomb sandwich structure based on the stress state data, and determine a structural design evaluation parameter of the initial laminate honeycomb sandwich structure under each of the failure modes based on the stress state data;

[0016] A judgment unit is used to judge whether the initial laminate honeycomb sandwich structure meets the design requirements based on each of the structural design evaluation parameters and the structural property data; if the initial laminate honeycomb sandwich structure meets the design requirements, the initial laminate honeycomb sandwich structure is determined as a frozen laminate honeycomb sandwich structure applied to an aircraft fuselage structure; otherwise, the structural parameters of the initial laminate honeycomb sandwich structure are optimized.

[0017] According to a third aspect of the present invention, there is provided a computer-readable storage medium having a computer program stored thereon, which, when executed by a processor, implements the above-mentioned design method for the honeycomb sandwich structure of aircraft laminates.

[0018] According to a fourth aspect of the present invention, there is provided a computer device comprising a memory, a processor and a computer program stored in the memory and executable on the processor, wherein the processor implements the above-mentioned design method for aircraft laminate honeycomb sandwich structures when executing the program.

[0019] According to a design method, device and storage medium of an aircraft laminate honeycomb sandwich structure provided by the present invention, compared with the current method of directly subjecting the aircraft laminate honeycomb sandwich structure designed according to human experience to static testing and optimizing the design of the aircraft laminate honeycomb sandwich structure according to the static test results, the present invention constructs a natural grid finite element model of the aircraft fuselage structure with an initial laminate honeycomb sandwich structure, and applies the same simulated loads, simulated displacement constraints and simulated environmental conditions as those borne by the aircraft during actual flight service to the natural grid finite element model, and simulates the aircraft service process based on the above-mentioned simulated loads, simulated displacement constraints and simulated environmental conditions, and determines the stress state of the initial laminate honeycomb sandwich structure after applying the simulated loads, simulated displacement constraints and simulated flight environmental conditions based on the simulation results, determines the failure mode according to the stress state, and calculates the evaluation parameters under each failure mode, and finally calculates the evaluation parameters under each failure mode and the attribute data of the initial laminate honeycomb sandwich structure according to the evaluation parameters under each failure mode. According to the invention, the structure of the initial laminate honeycomb sandwich structure is optimized, and thus, in the process of optimizing the design of the laminate honeycomb sandwich structure, the laminate honeycomb sandwich structure is substituted into the overall structure of the aircraft for stress state analysis, which can not only analyze the stress state of the laminate honeycomb sandwich structure in the external environment alone, but also analyze the structural stress state between the laminate honeycomb sandwich structure and other structures of the aircraft fuselage, so that the stress state analysis of the laminate honeycomb sandwich structure is more comprehensive, thereby improving the design accuracy of the laminate honeycomb sandwich structure, and at the same time, by simulating the actual flight state of the aircraft to obtain the stress state of the laminate honeycomb sandwich structure, the complex and changeable service environment of the aircraft can be considered, thereby ensuring that the stress state obtained by the analysis is more in line with the actual aircraft conditions, thereby further improving the design accuracy of the laminate honeycomb sandwich structure, and further, the present invention comprehensively considers various failure modes related to the design in the optimization design of the laminate honeycomb sandwich structure, thereby improving the design accuracy of the laminate honeycomb sandwich structure. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] The drawings described herein are used to provide a further understanding of the present invention and constitute a part of this application. The exemplary embodiments of the present invention and their descriptions are used to explain the present invention and do not constitute an improper limitation of the present invention. In the drawings:

[0021] Figure 1 A flow chart of a design method for a honeycomb sandwich structure of an aircraft laminate provided by an embodiment of the present invention is shown;

[0022] Figure 2 A flow chart of another design method of a honeycomb sandwich structure of an aircraft laminate provided by an embodiment of the present invention is shown;

[0023] Figure 3A schematic structural diagram of an initial aircraft laminate honeycomb sandwich structure provided by an embodiment of the present invention is shown;

[0024] Figure 4 A schematic diagram of the overall coordinate system of the entire aircraft (or fuselage) to which an aircraft laminate honeycomb sandwich structure provided by an embodiment of the present invention is applied is shown;

[0025] Figure 5 Schematic diagrams showing various failure modes of the honeycomb sandwich structure of an aircraft laminate provided by an embodiment of the present invention;

[0026] Figure 6 A schematic structural diagram of a design device for a honeycomb sandwich structure of an aircraft laminate provided by an embodiment of the present invention is shown;

[0027] Figure 7 A schematic diagram of the physical structure of a computer device provided by an embodiment of the present invention is shown. DETAILED DESCRIPTION

[0028] The present invention will be described in detail below with reference to the accompanying drawings and in combination with embodiments. It should be noted that the embodiments and features in the embodiments of the present application can be combined with each other without conflict.

[0029] At present, the aircraft laminate honeycomb sandwich structure designed based on human experience is directly subjected to static testing alone, and the aircraft laminate honeycomb sandwich structure is optimized according to the static test results, resulting in low accuracy in the design of the aircraft laminate honeycomb sandwich structure.

[0030] In order to solve the above problems, an embodiment of the present invention provides a design method for an aircraft laminate honeycomb sandwich structure, such as Figure 1 As shown, the method includes:

[0031] 101. Determine the initial laminate honeycomb sandwich structure of the laminate honeycomb sandwich to be designed, obtain the fuselage property data of the aircraft fuselage structure with the initial laminate honeycomb sandwich structure, and construct a natural grid finite element model of the aircraft fuselage structure based on the fuselage property data, wherein the fuselage property data includes the structural property data of the initial laminate honeycomb sandwich structure.

[0032] Among them, the fuselage attribute data include the three-dimensional structural dimension data of the aircraft fuselage structure, the material type and material properties of each component in the aircraft fuselage structure (such as the elastic modulus, Poisson's ratio, density, yield strength and other mechanical performance parameters of each component), etc., and the structural attribute data include the three-dimensional structural dimension data, material type, material properties (such as elastic modulus, Poisson's ratio, density, yield strength and other mechanical performance parameters) of the initial laminate honeycomb sandwich structure, etc.

[0033] Specifically, firstly, according to the initial structural design requirements of the laminate and the initial structural design requirements of the honeycomb sandwich to be designed, an initial laminate honeycomb sandwich structure is designed, and then the three-dimensional model of the initial laminate honeycomb sandwich structure is applied to the corresponding position of the aircraft body to obtain the three-dimensional digital model of the aircraft, and then based on the three-dimensional digital model of the aircraft, a three-dimensional finite element model of the aircraft is constructed using finite element software, that is, a natural grid finite element model of the aircraft body structure, and then according to the natural grid finite element model of the aircraft body structure, the stress state information of the initial laminate honeycomb sandwich structure during the actual flight is analyzed, and the initial laminate honeycomb sandwich structure is continuously optimized based on the accepted state information, so as to obtain a laminate honeycomb sandwich structure whose strength, stiffness and weight all meet the design requirements. Therefore, in the process of optimizing the design of the laminate honeycomb sandwich structure, the laminate honeycomb sandwich structure is substituted into the overall structure of the aircraft for stress state analysis, which can not only analyze the stress state of the laminate honeycomb sandwich structure in the external environment alone, but also analyze the structural stress state between the laminate honeycomb sandwich structure and other structures of the aircraft fuselage, making the stress state analysis of the laminate honeycomb sandwich structure more comprehensive. At the same time, the three-dimensional finite element model of the aircraft can be used to accurately analyze the laminate honeycomb sandwich structure and predict its response characteristics under various working conditions, thereby improving the design accuracy. The finite element model supports rapid iterative calculations, and multiple design schemes can be evaluated and compared in a short time, thereby improving design efficiency and reducing costs.

[0034] 102. Determine the actual loads, actual displacement constraints, and actual flight environment conditions that the actual aircraft will bear during its service; determine simulated loads that are the same as the actual loads, simulated displacement constraints that are the same as the actual displacement constraints, and simulated flight environment conditions that are the same as the actual flight environment conditions.

[0035] Among them, the simulated loads include aerodynamic loads, inertial loads, concentrated loads, etc.; the simulated displacement constraints include: node displacement constraints, aircraft front landing gear displacement constraints, aircraft rear landing gear displacement constraints, etc.; the simulated flight environment conditions include humidity conditions, temperature conditions, etc.

[0036] Specifically, by collecting historical data, it is possible to determine the actual loads such as aerodynamic loads, inertial loads, concentrated loads, etc. of the aircraft used in the designed laminate honeycomb sandwich structure during the actual flight service, as well as the actual displacement constraints such as node displacement constraints, aircraft front landing gear displacement constraints, aircraft rear landing gear displacement constraints, and humidity conditions, temperature conditions, and other actual flight environment conditions, and determine the actual loads, actual displacement constraints, and actual flight environment conditions as the simulated loads, simulated displacement constraints, and simulated flight environment conditions during the service of the aircraft simulation aircraft, and then use the above-mentioned simulated loads, simulated displacement constraints, simulated flight environment conditions, and the natural grid finite element model of the aircraft body structure to simulate the flight service of the aircraft, and finally optimize the structural design of the initial laminate honeycomb sandwich structure according to the simulation results. Thus, by determining the actual loads, actual displacement constraints, and actual flight environment conditions, it is possible to ensure that the simulated flight process of the aircraft is closer to the actual flight service process, so that the stress state analysis of the initial laminate honeycomb sandwich structure can be more accurate, and then the optimization design accuracy of the laminate honeycomb sandwich structure can be improved.

[0037] 103. Apply simulated loads to the natural grid finite element model of the aircraft body structure, set simulated displacement constraints, set simulated flight environment conditions, and determine the stress state data of the initial laminate honeycomb sandwich structure after applying the simulated loads, setting the simulated displacement constraints, and setting the simulated flight environment conditions.

[0038] Among them, the stress state data include: the displacement at the center of the skin of the initial laminate honeycomb sandwich structure, the stress, strength, strain data, and structural dimensions of the honeycomb core in all directions, the stress, structural dimensions of the laminate in all directions, the material property data of the honeycomb sandwich and the laminate, etc.; the stress state data also includes: load type, load size, load direction, load distribution information, etc.

[0039] Specifically, simulated loads are applied to the natural grid finite element model of the aircraft body structure, simulated displacement constraints are set, and simulated flight environment conditions are set. Based on the model, simulated loads are applied, simulated displacement constraints are set, and simulated flight environment conditions are set. The natural grid finite element model of the aircraft body structure is used to simulate the aircraft service process. In the simulation process, the stress state data of the initial laminate honeycomb sandwich structure after applying the simulated loads, setting the simulated displacement constraints, and setting the simulated flight environment conditions are determined. Finally, the initial laminate honeycomb sandwich structure is optimized according to the stress state data. In this way, the laminate honeycomb sandwich structure is substituted into the entire aircraft during the process of optimizing the laminate honeycomb sandwich structure. The stress state analysis of the body structure can not only analyze the stress state of the external environment of the laminate honeycomb sandwich structure alone, but also analyze the structural stress state between the laminate honeycomb sandwich structure and other structures of the aircraft fuselage, making the stress state analysis of the laminate honeycomb sandwich structure more comprehensive, thereby improving the design accuracy of the laminate honeycomb sandwich structure. At the same time, in the simulation process, by applying simulated loads, setting simulated displacement constraints, and setting simulated flight environment conditions, various constraints in the actual service process of the aircraft are imposed on the natural grid finite element model of the aircraft fuselage structure, which can ensure the accuracy of the aircraft flight simulation, thereby improving the analysis accuracy of the stress state data of the initial laminate honeycomb sandwich structure.

[0040] 104. Based on the stress state data, determine the various failure modes of the initial laminate honeycomb sandwich structure, and based on the stress state data, determine the structural design evaluation parameters of the initial laminate honeycomb sandwich structure under each failure mode.

[0041] Among them, the various failure modes include: honeycomb core crushing failure mode, honeycomb core shear tearing failure mode, laminate destruction failure mode, lattice buckling failure mode, panel wrinkling failure mode, shear wrinkle failure mode, overall buckling failure mode, etc.; the structural design evaluation parameters can be the safety margin under each failure mode.

[0042] Specifically, according to the load type, load size, load direction, load distribution information, etc. borne by the initial laminate honeycomb sandwich structure, the honeycomb core crushing failure mode, honeycomb core shear tearing failure mode, laminate destruction failure mode, lattice buckling failure mode, panel wrinkling failure mode, shear wrinkle failure mode, overall buckling failure mode, etc. of the initial laminate honeycomb sandwich structure are determined. Then, according to the displacement at the center of the skin of the initial laminate honeycomb sandwich structure, the stress, strength, strain data of the honeycomb core in all directions, the structural dimensions, the stress, structural dimensions of the laminate in all directions, the material property data of the honeycomb sandwich and the laminate, etc., the structural design evaluation parameters of the initial laminate honeycomb sandwich structure under each failure mode are calculated, and the initial laminate honeycomb sandwich structure is optimized according to the structural design evaluation parameters. Therefore, through a comprehensive analysis of various failure modes, it can be clarified under what circumstances the initial laminate honeycomb sandwich structure is prone to damage, thereby providing an accurate basis for the optimization design of the initial laminate honeycomb sandwich structure, and further improving the design accuracy of the laminate honeycomb sandwich structure.

[0043] 105. Based on various structural design evaluation parameters and structural property data, determine whether the initial laminate honeycomb sandwich structure meets the design requirements. If the initial laminate honeycomb sandwich structure meets the design requirements, the initial laminate honeycomb sandwich structure is determined as a frozen laminate honeycomb sandwich structure applied to the aircraft fuselage structure. Otherwise, the structural parameters of the initial laminate honeycomb sandwich structure are optimized.

[0044] Among them, design requirements include stiffness design requirements, strength design requirements, weight design requirements, economic design requirements, durability design requirements, bonding strength design requirements, stability design requirements, etc.

[0045] Specifically, according to various structural design evaluation parameters and structural property data, it is judged whether the initial laminate honeycomb sandwich structure meets the stiffness design requirements, strength design requirements, weight design requirements, economic design requirements, durability design requirements, bonding strength design requirements, stability design requirements, etc. If only the stiffness design requirements, strength design requirements, and weight design requirements are focused on, then when the initial laminate honeycomb sandwich structure meets the stiffness design requirements, strength design requirements, and weight design requirements at the same time, the initial laminate honeycomb sandwich structure is determined to be the optimal structure, that is, the structure that meets the design requirements. At this time, there is no need to structurally optimize the initial laminate honeycomb sandwich structure, and the initial laminate honeycomb sandwich structure can be directly applied to the design of the aircraft. If the initial laminate honeycomb sandwich structure does not meet any of the stiffness design requirements, strength design requirements, and weight design requirements, it is necessary to structurally optimize the initial laminate honeycomb sandwich structure until the laminate honeycomb sandwich structure after the optimized structure meets various design requirements. The embodiment of the present invention comprehensively considers various failure modes related to the design in the optimization design of the laminated honeycomb sandwich structure, thereby improving the design accuracy of the laminated honeycomb sandwich structure.

[0046] According to a design method of an aircraft laminate honeycomb sandwich structure provided by the present invention, compared with the prior art, in the embodiment of the present invention, by substituting the laminate honeycomb sandwich structure into the overall structure of the aircraft for stress state analysis during the process of optimizing the design of the laminate honeycomb sandwich structure, not only the stress state of the laminate honeycomb sandwich structure in the external environment alone can be analyzed, but also the structural stress state between the laminate honeycomb sandwich structure and other structures of the aircraft fuselage can be analyzed, so that the stress state analysis of the laminate honeycomb sandwich structure is more comprehensive, thereby improving the design accuracy of the laminate honeycomb sandwich structure. At the same time, by simulating the actual flight state of the aircraft to obtain the stress state of the laminate honeycomb sandwich structure, the complex and changeable service environment of the aircraft can be considered, thereby ensuring that the stress state obtained by analysis is more in line with the actual aircraft conditions, thereby further improving the design accuracy of the laminate honeycomb sandwich structure. Furthermore, in the optimization design of the laminate honeycomb sandwich structure, the present invention comprehensively considers various failure modes related to the design, thereby improving the design accuracy of the laminate honeycomb sandwich structure.

[0047] Further, in order to better illustrate the above process of classifying data, as a refinement and extension of the above embodiment, the embodiment of the present invention provides another design method for an aircraft laminate honeycomb sandwich structure, such as Figure 2 As shown, the method includes:

[0048] 201. Determine an initial laminate honeycomb sandwich structure of the laminate honeycomb sandwich to be designed, obtain fuselage property data of an aircraft fuselage structure with the initial laminate honeycomb sandwich structure, and construct a natural grid finite element model of the aircraft fuselage structure based on the fuselage property data, wherein the fuselage property data includes structural property data of the initial laminate honeycomb sandwich structure.

[0049] For the embodiment of the present invention, in order to optimize the design of the laminate honeycomb sandwich structure, it is first necessary to design an initial laminate honeycomb sandwich structure. Based on this, step 201 specifically includes: determining the laminate initial structure design requirements of the laminate to be designed in the laminate honeycomb sandwich to be designed and the honeycomb sandwich initial structure design requirements of the honeycomb sandwich to be designed, wherein the laminate initial structure design requirements include: at least two fiber directions are arranged, the ply contacting the core and the surface ply need to be continuous, an anti-penetration film is laid, the layer is lost in the core area, and the first layer is lost The distance from the honeycomb edge is greater than the first preset distance The minimum distance between two removed layers is greater than a second preset distance threshold, and the initial structural design requirements of the honeycomb sandwich core include: the direction of the honeycomb core band is parallel to the main load direction, the radius of the honeycomb corner area is greater than a preset radius threshold, the honeycomb bevel angle is a preset angle threshold, the height of the honeycomb bevel edge is greater than a preset height threshold, a laminate liner is embedded in the honeycomb core, and adhesives and foam cores of preset materials are used on the inside and outside of the honeycomb core; based on the initial structural design requirements of the laminate and the initial structural design requirements of the honeycomb sandwich to be designed, the initial laminate honeycomb sandwich structure is determined.

[0050] Among them, the first preset distance threshold, the second preset distance threshold, the preset radius threshold, the preset angle threshold, the preset height threshold, and the preset material are all set according to actual needs.

[0051] Specifically, Figure 3 As shown, for example, for the laminate to be designed in the honeycomb sandwich of the laminate to be designed, (1) at least two fiber directions are arranged; (2) the ply in contact with the core and the surface ply must be continuous; (3) an anti-permeability film (Tedlar) is laid; (4) the dropped layers are in the core area, and the first dropped layer is at least 10 mm away from the honeycomb edge; (5) the minimum distance between two dropped layers is greater than 5 mm. For the honeycomb sandwich of the laminate to be designed: (1) the core belt direction is parallel to the main load direction; (2) the corner area should use a minimum radius of 20 mm; (3) the honeycomb bevel angle is 20°±5°, and the minimum bevel edge height is 1.5 mm; (4) if an opening is required, a laminate liner is embedded in the honeycomb core, and a foaming adhesive and a foam core are used on its inner and outer sides. Based on the above information, the initial laminate honeycomb sandwich structure can be designed.

[0052] Furthermore, the fuselage attribute data includes: fuselage structure size data and fuselage material attribute data of the aircraft fuselage structure with an initial laminate honeycomb sandwich structure. After the initial laminate honeycomb sandwich structure is designed, a natural mesh finite element model of the aircraft fuselage structure is constructed based on the three-dimensional digital model of the aircraft with the initial laminate honeycomb sandwich structure. Based on this, step 201 also includes: determining the coordinate system information, unit information, and mesh division strategy information required to construct the natural mesh finite element model of the aircraft fuselage structure, wherein the coordinate system information is to select the overall coordinate system of the aircraft fuselage structure, define a spatial point on the aircraft symmetry plane directly in front of the aircraft nose as the origin O in the overall coordinate system, the x-axis in the overall coordinate system points to the left side of the aircraft along the heading, the x-axis is perpendicular to the aircraft fuselage along the heading, the y-axis in the overall coordinate system passes through the center of the arc of the typical cross-section of the aircraft fuselage, the y-axis points to the opposite side of the aircraft along the heading, and the z The axis follows the right-hand rule and passes through the origin O vertically upward. The meshing strategy information includes mesh nodes, unit numbers, unit attributes, and mesh sizes. The mesh nodes are selected at the intersections of the theoretical outer shape surface of the skin of the aircraft body structure, the theoretical outer shape plane of the beam, and the mid-surface of the long stringer, the mid-surface of the rib, and the mid-surface of the frame. The unit numbering adopts beam elements, rod elements, shell elements, supplemented by rigid units, constraint units, scalar units, and three-dimensional volume elements. The initial laminate honeycomb sandwich structure adopts shell element discretization, and the initial laminate honeycomb sandwich structure adopts a plying method to define unit attributes. Based on the body structure size data, body material attribute data, coordinate system information, unit information, and meshing strategy information, a natural mesh finite element model of the aircraft body structure is constructed using a preset finite element modeling software.

[0053] Specifically, the natural mesh finite element model of the aircraft body structure can be constructed using modeling software such as MSC.Patran and Hypermesh. In addition to the three-dimensional digital model of the aircraft, the required parameters include: modeling standards, aircraft structure dimensions, materials and performance of each aircraft component, meshing strategy, etc. The modeling process includes: coordinate system definition, unit definition, material property definition, meshing, node and unit numbering, unit property definition and model checking. The specific modeling process is: First, the natural mesh finite element model selects the overall coordinate system of the entire aircraft (or fuselage), such as Figure 4As shown in the figure, the overall coordinate system of the whole aircraft defines a space point on the symmetry plane (yz plane) in front of the nose as the origin O, the x-axis points to the left side of the aircraft along the heading, the y-axis passes through the center of the arc of the typical section of the fuselage, points from the nose to the tail, and the z-axis follows the right-hand rule and points vertically upward. The unit system of the natural mesh finite element model is as follows: the unit of length is millimeter (mm), the unit of mass is ton (t), the unit of force is Newton (N), the unit of temperature is degree Celsius (℃), the unit of time is second (s), and the unit of stress is megapascal (MPa). In the natural mesh finite element model, both the laminate and the honeycomb are defined as orthotropic material properties, and the material properties that need to be input include: elastic modulus, shear modulus, Poisson's ratio, density, and thermal expansion coefficient. The mesh nodes of the natural mesh finite element model are selected at the intersection of the theoretical outer shape surface of the skin, the theoretical outer shape plane of the beam, and the mid-surface of the long stringer with the mid-surface of the rib and the mid-surface of the frame. The elements mainly include beam element (CBAR, Beam Element), rod element (CROD, Rod Element), shell element (CQUAD4 / CTRIA3), supplemented by rigid element, constraint element, scalar element and three-dimensional element. The laminate / honeycomb sandwich structure is discretized by shell element (CQUAD4 / CTRIA3). The laminate / honeycomb sandwich structure uses a layering method to define the unit properties and gives the material true thickness and direction. Model checks include: unit system, geometric dimensions, free boundaries, unit normals, repeated nodes, repeated units, unit mass, material properties, material coordinate system, modal and ground calculation checks. MSC.Nastran is used to solve the natural mesh finite element model, and based on the results, the internal load of the structure is extracted. Therefore, the natural mesh finite element model of the aircraft fuselage structure can be constructed according to the above method.

[0054] 202. Determine the actual loads, actual displacement constraints, and actual flight environment conditions that the actual aircraft will bear during its service; determine simulated loads that are the same as the actual loads, simulated displacement constraints that are the same as the actual displacement constraints, and simulated flight environment conditions that are the same as the actual flight environment conditions.

[0055] 203. Apply simulated loads to the natural grid finite element model of the aircraft body structure, set simulated displacement constraints, set simulated flight environment conditions, and determine the stress state data of the initial laminate honeycomb sandwich structure after applying the simulated loads, setting the simulated displacement constraints, and setting the simulated flight environment conditions.

[0056] Among them, the simulated loads include: aerodynamic loads, inertial loads, concentrated loads, the simulated displacement constraints include: node displacement constraints, aircraft front landing gear displacement constraints, aircraft rear landing gear displacement constraints, and the simulated flight environment conditions include humidity conditions and temperature conditions.

[0057] For the embodiment of the present invention, the actual load, actual displacement constraint condition, and actual flight environment condition borne by the actual aircraft during service are determined as simulated load, simulated displacement constraint condition, and simulated flight environment condition, and based on the simulated load, simulated displacement constraint condition, simulated flight environment condition, and the natural grid finite element model of the aircraft body structure, the simulation of the aircraft flight service is realized, and the stress state data of the initial laminate honeycomb sandwich structure is determined through the simulation result. Based on this, step 203 specifically includes: applying the aerodynamic load and the inertial load to the rib, frame, and beam station surface skin nodes of the natural grid finite element model of the aircraft body structure respectively The concentrated load is applied to the corresponding structural unit through the rigid body unit; any three nodes of the aircraft are selected in the natural grid finite element model of the aircraft body structure to apply node displacement constraints, the displacement constraints of the aircraft front landing gear in the x-axis direction and the z-axis direction in the overall coordinate system are applied to the front landing gear of the aircraft in the natural grid finite element model of the aircraft body structure, and the displacement constraints of the aircraft rear landing gear in the x-axis, y-axis and z-axis directions in the overall coordinate system are respectively applied to the two rear landing gears of the aircraft in the natural grid finite element model of the aircraft body structure; the humidity condition and the temperature condition are set for the natural grid finite element model of the aircraft body structure.

[0058] Specifically, based on the established natural mesh finite element model of the aircraft fuselage structure, the loads during the service life of the aircraft are applied to it, and the displacement constraints and environmental constraints are simulated. The specific method is: the aerodynamic load and inertial load are applied to the rib, frame, and beam surface skin nodes of the aircraft fuselage structure in the natural mesh finite element model, and the concentrated load is applied to the corresponding structural unit through the rigid body unit (RBE2). In the natural mesh finite element model, any three nodes of the aircraft are selected to apply node displacement constraints, and the following are applied at the front landing gear: Figure 4 The displacement constraints of the aircraft's front landing gear in the x-axis and z-axis directions of the coordinate system are applied respectively at the two rear landing gears as follows Figure 4 The displacement constraint of the aircraft rear landing gear in the z-axis direction and the displacement constraint conditions of the aircraft rear landing gear in the x, y and z-axis directions in the coordinate system. At the same time, humidity constraints and temperature constraints are set for the natural grid finite element model of the aircraft body structure. Therefore, various loads and constraints in the actual service process of the aircraft are applied to the natural grid finite element model of the aircraft body structure, which can make the aircraft service simulation more accurate, thereby improving the design accuracy of the laminate honeycomb sandwich structure, and making the designed laminate honeycomb sandwich structure more in line with actual needs.

[0059] 204. Based on the stress state data, various failure modes of the initial laminate honeycomb sandwich structure are determined, and based on the stress state data, structural design evaluation parameters of the initial laminate honeycomb sandwich structure under various failure modes are determined.

[0060] Among them, Figure 5As shown, the various failure modes include honeycomb core crushing failure mode, honeycomb core shear tearing failure mode, laminate damage failure mode, lattice buckling failure mode, panel wrinkling failure mode, shear wrinkle failure mode, and overall buckling failure mode. For the embodiment of the present invention, based on the load type, load size, load direction, load distribution information, etc. borne by the initial laminate honeycomb sandwich structure in the stress state data, the honeycomb core crushing failure mode, honeycomb core shear tearing failure mode, laminate damage failure mode, lattice buckling failure mode, panel wrinkling failure mode, shear wrinkle failure mode, and overall buckling failure mode of the initial laminate honeycomb sandwich structure are determined. Then, in order to determine whether the initial laminate honeycomb sandwich structure meets the design requirements, it is also necessary to determine the structural design evaluation parameters of the initial laminate honeycomb sandwich structure under each failure mode. Based on this, step 20 4 specifically includes: determining the honeycomb core crushing evaluation parameters under the honeycomb core crushing failure mode based on the stress in the thickness direction of the honeycomb core in the honeycomb sandwich structure of the initial laminate and the compressive strength of the honeycomb sandwich material; determining the honeycomb core shear tearing evaluation parameters under the honeycomb core shear tearing failure mode based on the longitudinal shear stress and transverse shear stress of the honeycomb sandwich in the honeycomb sandwich structure of the initial laminate, and the longitudinal shear strength and transverse shear strength of the honeycomb sandwich material; determining the honeycomb core shear tearing evaluation parameters under the honeycomb core shear tearing failure mode based on the x-axis stress in the x-axis direction, the y-axis stress in the y-axis direction, and the x-axis stress in the y-axis direction of the laminate in the honeycomb sandwich structure of the initial laminate respectively in the overall coordinate system. and y-axis directions, the x-axis shear stress corresponding to the x-axis and y-axis directions, the x-axis allowable stress in the x-axis direction of the laminate corresponding to the prepreg in the overall coordinate system, the y-axis allowable stress in the y-axis direction, and the xy-axis shear allowable stress corresponding to the x-axis and y-axis directions, determine the laminate failure evaluation parameters under the laminate failure mode; based on the Poisson's ratio of the laminate in the initial laminate honeycomb sandwich structure, the x-axis compressive elastic modulus of the laminate along the x-axis direction, the y-axis compressive elastic modulus along the y-axis direction, the shear modulus of the laminate in the xy plane where the x-axis and y-axis are located, the laminate thickness of the laminate, The honeycomb inscribed circle diameter of the honeycomb sandwich structure in the initial laminate honeycomb sandwich structure, the x-axis stress of the honeycomb wall panel unit of the honeycomb sandwich structure in the x-axis direction, the y-axis stress in the y-axis direction, the lattice buckling shear design allowable value of the honeycomb sandwich wall panel, the shear stress of the honeycomb wall panel unit, the first calculation parameter, and the second calculation parameter are used to determine the lattice buckling evaluation parameters under the lattice buckling failure mode, wherein the lattice buckling evaluation parameters include: the lattice buckling evaluation parameters under the compression load, the lattice buckling evaluation parameters under the shear load, and the lattice buckling evaluation parameters under the combined action of the compression load and the shear load;Based on the length of the laminate in the load direction, the thickness of the laminate, the bending stiffness coefficient of the laminate, the half-wave number in the load direction, the width of the laminate perpendicular to the load direction, the panel centroid spacing of the laminate, the z-axis compressive elastic modulus of the laminate along the z-axis direction, the x-axis stress of the honeycomb wall panel unit in the x-axis direction, and the y-axis stress in the y-axis direction of the honeycomb sandwich structure, the panel wrinkling evaluation parameters under the panel wrinkling failure mode are determined; based on the initial laminate honeycomb sandwich structure, the panel wrinkling evaluation parameters under the panel wrinkling failure mode are determined; based on the initial laminate honeycomb sandwich structure, the panel wrinkling evaluation parameters under the panel wrinkling failure mode are determined. The honeycomb thickness of the honeycomb sandwich structure in the core structure, the laminate thickness of different laminates, the equivalent elastic modulus of different laminates, the shear modulus of the honeycomb sandwich in the xy plane where the x-axis and y-axis are located, the x-axis stress of the honeycomb wall plate unit of the honeycomb sandwich structure in the x-axis direction, and the y-axis stress in the y-axis direction, determine the shear wrinkle evaluation parameters under the shear wrinkle failure mode; based on the critical buckling load and actual load of the honeycomb sandwich structure of the initial laminate, determine the overall buckling evaluation parameters under the overall buckling failure mode. ;

[0061] Specifically, the honeycomb core crush evaluation parameter is calculated according to the following formula:

[0062]

[0063] Among them, MS a is the honeycomb core crushing evaluation parameter, in MPa, [σ ac ] is the compressive strength of the honeycomb sandwich material, σ az is the stress in the thickness direction of the honeycomb core.

[0064] The honeycomb core shear tearing evaluation parameters are calculated according to the following formula:

[0065]

[0066] Among them, MS b is the honeycomb core shear tearing evaluation parameter, τ LT is the longitudinal shear stress of the honeycomb core, [τ LT ] is the longitudinal shear strength of the honeycomb sandwich material, τ WT is the transverse shear stress of the honeycomb core, [τ WT ] is the transverse shear strength of the honeycomb core.

[0067] The laminate damage evaluation parameters are calculated according to the following formula:

[0068]

[0069] Among them, F1 is the failure strength of the laminate, σ x is the x-axis stress of the laminates in the initial laminate honeycomb sandwich structure in the x-axis direction in the global coordinate system, σ yis the y-axis stress of the laminates in the initial laminate honeycomb sandwich structure in the y-axis direction in the global coordinate system, σ xy is the xy-axis shear stress corresponding to the x-axis and y-axis directions of the laminate in the overall coordinate system, X is the x-axis allowable stress of the laminate corresponding to the prepreg in the x-axis direction of the overall coordinate system, Y is the y-axis allowable stress of the laminate corresponding to the prepreg in the y-axis direction of the overall coordinate system, S is the xy-axis shear stress corresponding to the x-axis and y-axis directions of the laminate corresponding to the prepreg in the overall coordinate system, MS c Evaluation parameters for laminate failure.

[0070] Under the compressive load, the lattice buckling evaluation parameter is calculated according to the following formula:

[0071]

[0072]

[0073] Among them, [σ cr,x ] is the allowable value of the buckling design of the sandwich panel along the x-axis in the honeycomb sandwich structure of the initial laminate, [σ cr,y ] is the allowable value of the buckling design of the sandwich panel along the y-axis in the honeycomb sandwich structure of the initial laminate, μ x is the Poisson’s ratio of the laminate along the x-axis in the initial laminate honeycomb sandwich structure, μ y is the Poisson’s ratio of the laminate along the y-axis in the initial laminate honeycomb sandwich structure, E fx is the x-axis compressive elastic modulus of the laminate along the x-axis direction, E fy is the y-axis compressive elastic modulus of the laminate along the y-axis direction, G xy is the shear modulus of the laminate in the xy plane where the x-axis and y-axis are located, t is the thickness of the laminate, s is the diameter of the honeycomb inscribed circle, σ p,x is the x-axis stress of the honeycomb wall panel unit in the x-axis direction, σ p,y is the y-axis stress of the honeycomb wall unit in the y-axis direction, MS d Lattice buckling evaluation parameters for compressive loading

[0074] Under the action of shear load, the lattice buckling evaluation parameters are calculated according to the following functional formula:

[0075]

[0076] Among them, MS e is the lattice buckling evaluation parameter under shear load, [τ cr ] is the allowable value of sandwich panel lattice buckling shear design, τ is the shear stress of honeycomb panel unit, E fxis the x-axis compressive elastic modulus of the laminate along the x-axis direction, E fy is the y-axis compressive elastic modulus of the laminate along the y-axis direction, t is the thickness of the laminate, and s is the diameter of the honeycomb inscribed circle.

[0077] When the sandwich structure is subjected to compression and shear loads at the same time, the lattice buckling evaluation parameter is calculated according to the following formula:

[0078]

[0079] or

[0080]

[0081] Among them, MS f is a lattice buckling evaluation parameter when subjected to compression and shear loads at the same time, a is a first calculation parameter, which can be a first safety margin calculation parameter. For example, the first calculation parameter can be set between 1.95 and 2.05 according to actual needs, b is a second calculation parameter, which can be a second safety margin calculation parameter. For example, the second calculation parameter can be set between 3.98 and 4.03 according to actual needs, [σ cr,x ] is the allowable value of the buckling design of the sandwich panel along the x-axis in the honeycomb sandwich structure of the initial laminate, [σ cr,y ] is the allowable value of the buckling design of the sandwich panel along the y-axis in the honeycomb sandwich structure of the initial laminate, σ p,x is the x-axis stress of the honeycomb wall panel unit in the x-axis direction, σ p,y is the y-axis stress of the honeycomb wall panel unit in the y-axis direction, [τ cr ] is the allowable value of lattice buckling shear design for sandwich panel, and τ is the shear stress of honeycomb panel unit.

[0082] In summary, the embodiment of the present invention calculates the lattice buckling evaluation parameters by comprehensively considering various influencing factors affecting the lattice buckling evaluation parameters, and introduces calculation parameters, thereby improving the calculation accuracy of the lattice buckling evaluation parameters. The embodiment of the present invention calculates the lattice buckling evaluation parameters under various loads, and designs the laminate honeycomb sandwich structure through the evaluation parameters under various loads, thereby improving the design accuracy of the laminate honeycomb sandwich structure.

[0083] The panel wrinkling evaluation parameters are calculated according to the following formula:

[0084] [σ ccr,x ]or

[0085] or

[0086] Among them, MS g is the panel wrinkling evaluation parameter, [σ ccr,x ] is the allowable value of laminate panel wrinkling design, [σ ccr,y ] is the design allowable value of wrinkling of honeycomb sandwich panel, L is the length of laminate in the loading direction, t is the thickness of laminate panel, D 11 , D 12 , D 66 , D 22 are the bending stiffness coefficients of the laminate, m is the half-wave number in the loading direction (L direction), b is the width of the laminate perpendicular to the loading direction, h is the distance between the panel centroids of the laminate, E cz is the z-axis compressive elastic modulus of the laminate along the z-axis direction, σ p,x is the x-axis stress of the honeycomb wall panel unit in the x-axis direction, σ p,y is the y-axis stress of the honeycomb wall panel unit in the y-axis direction.

[0087] Under the action of compression load, the shear wrinkle evaluation parameters are calculated according to the following formula:

[0088]

[0089] or

[0090] Among them, MS h is the shear wrinkle evaluation parameter under compression load, [σ cr1,x ] is the design allowable value of the first plywood under compression load, [σ cr1,y ] is the design allowable value of the first honeycomb core under compression load, [σ cr2,x ] is the design allowable value of the second laminate under compression load, [σ cr2,y ] is the design allowable value of the second honeycomb core under compression load, t c is the honeycomb thickness, t1 is the thickness of the first laminate, t2 is the thickness of the second laminate, E1 is the equivalent elastic modulus of the first laminate, E2 is the equivalent elastic modulus of the second laminate, G ij is the shear modulus of the honeycomb core in the xy plane where the x-axis and y-axis are located, σ p,x is the x-axis stress of the honeycomb wall panel unit in the x-axis direction, σ p,y is the y-axis stress of the honeycomb wall panel unit in the y-axis direction.

[0091] Under the action of shear load, the shear wrinkle evaluation parameters are calculated according to the following formula:

[0092]

[0093] Among them, τ cr,1 is the design allowable value of shear wrinkling of the first ply, τ cr,2 is the design allowable value of shear wrinkling of the second plywood, t c is the honeycomb thickness, t1 is the thickness of the first laminate, t2 is the thickness of the second laminate, E1 is the equivalent elastic modulus of the first laminate, E2 is the equivalent elastic modulus of the second laminate, G xz is the shear modulus of the laminate when the load is parallel to the x-axis, G yz is the shear modulus of the laminate when the load is parallel to the y-axis, MS i is the shear wrinkle evaluation parameter under shear load, and τ is the shear stress of the honeycomb wall panel unit.

[0094] In summary, the embodiment of the present invention calculates the shear wrinkle evaluation parameters by comprehensively considering various influencing factors affecting the shear wrinkle evaluation parameters, thereby improving the calculation accuracy of the shear wrinkle evaluation parameters. The embodiment of the present invention calculates the shear wrinkle evaluation parameters under various loads, and designs the laminate honeycomb sandwich structure through the evaluation parameters under various loads, thereby improving the design accuracy of the laminate honeycomb sandwich structure.

[0095] The overall buckling evaluation parameters are calculated according to the following formula:

[0096]

[0097] Among them, MS j is the overall buckling evaluation parameter, [ν cr ] is the critical buckling load and ν is the actual bearing load.

[0098] The embodiments of the present invention respectively propose analytical calculation methods for working stress, allowable stress, and safety margin for each failure mode of the laminate / honeycomb sandwich structure, effectively avoiding the problems of large workload, non-convergence of iterations, and low calculation accuracy caused by finite element calculations. The failure modes of the laminate honeycomb sandwich structure under complex load conditions are given for the honeycomb core, honeycomb panel, and honeycomb siding locally and as a whole, which truly reflects the stress characteristics of the structure, ensures the integrity of the strength analysis, and can improve the design accuracy of the laminate honeycomb sandwich structure.

[0099] 205. Based on various structural design evaluation parameters, determine whether the initial laminate honeycomb sandwich structure meets the strength design requirements.

[0100] Among them, the preset parameter threshold is set according to actual needs. Specifically, after determining the structural design evaluation parameters under various failure modes corresponding to the initial laminate honeycomb sandwich structure, it is necessary to judge whether the initial laminate honeycomb sandwich structure meets the strength design requirements based on each structural design evaluation parameter. Based on this, step 205 specifically includes: judging whether the initial laminate honeycomb sandwich structure meets the strength design requirements based on the honeycomb core crushing evaluation parameter, honeycomb core shear tearing evaluation parameter, laminate damage evaluation parameter, lattice buckling evaluation parameter, panel wrinkling evaluation parameter, shear wrinkle evaluation parameter, and overall buckling evaluation parameter. The specific judgment method includes: judging whether the honeycomb core crushing evaluation parameter, honeycomb core shear tearing evaluation parameter, laminate damage evaluation parameter, lattice buckling evaluation parameter, panel wrinkling evaluation parameter, shear wrinkle evaluation parameter, and overall buckling evaluation parameter are all greater than the preset parameter threshold. If they are all greater than the preset parameter threshold, it is determined that the initial laminate honeycomb sandwich structure meets the strength design requirements.

[0101] Specifically, if the structural design parameters under each failure mode are greater than the preset parameter threshold, it is determined that the initial laminate honeycomb sandwich structure meets the strength design requirements. If any parameter among the structural design parameters under each failure mode is less than the preset parameter threshold, it is determined that the initial laminate honeycomb sandwich structure does not meet the strength design requirements. At this time, it is necessary to optimize the structural parameters of the initial laminate honeycomb sandwich structure until the optimized laminate honeycomb sandwich structure meets the strength design requirements.

[0102] 206. Based on the structural property data, determine whether the initial laminate honeycomb sandwich structure meets the stiffness design requirements and weight design requirements.

[0103] The structural attribute data include: honeycomb core thickness, honeycomb core shear modulus, bending stiffness, shear stiffness, structural weight, etc. corresponding to the initial laminate honeycomb sandwich structure.

[0104] For the embodiment of the present invention, when the initial laminate honeycomb sandwich structure meets the strength design requirements, it is also necessary to determine whether the initial laminate honeycomb sandwich structure meets the stiffness design requirements and the weight design requirements. Based on this, step 206 specifically includes: based on the honeycomb core thickness, honeycomb core shear modulus, bending stiffness, and shear stiffness, determining whether the initial laminate honeycomb sandwich structure meets the stiffness design requirements; based on the structure weight, determining whether the initial laminate honeycomb sandwich structure meets the weight design requirements.

[0105] Specifically, the necessary stiffness check is performed on the initial laminate honeycomb sandwich structure. The honeycomb core must have sufficient thickness and shear modulus, and the initial laminate honeycomb sandwich structure must have sufficient bending stiffness and shear stiffness to prevent excessive flexural deformation under the design load. If the initial laminate honeycomb sandwich structure does not meet the stiffness design requirements, the structural parameters of the initial laminate honeycomb sandwich structure are optimized until the optimized laminate honeycomb sandwich structure meets the strength design requirements and stiffness design requirements. On the basis of ensuring strength and stiffness, the weight index must also be considered, that is, the weight of the initial laminate honeycomb sandwich structure needs to be within a preset weight range (wherein the preset weight range is set according to actual needs). If the initial laminate honeycomb sandwich structure meets the strength design requirements, stiffness design requirements and weight design requirements at the same time, it is determined that the initial laminate honeycomb sandwich structure meets the design requirements. At this time, there is no need to optimize the initial laminate honeycomb sandwich structure, and it can be directly used as a structural design component of the aircraft.

[0106] 207. If the initial laminate honeycomb sandwich structure satisfies the strength design requirements, the stiffness design requirements, and the weight design requirements at the same time, then the initial laminate honeycomb sandwich structure is determined to meet the design requirements, and the initial laminate honeycomb sandwich structure is determined as a frozen laminate honeycomb sandwich structure applied to the aircraft fuselage structure. Otherwise, the structural parameters of the initial laminate honeycomb sandwich structure are optimized.

[0107] For an embodiment of the present invention, if the initial laminate honeycomb sandwich structure does not meet any one of the strength design requirements, the stiffness design requirements and the weight design requirements, it is necessary to optimize the structural parameters of the initial laminate honeycomb sandwich structure, and the specific optimization method includes: if the initial laminate honeycomb sandwich structure does not meet any one of the strength design requirements, the stiffness design requirements and the weight design requirements, it is determined that the initial laminate honeycomb sandwich structure does not meet the design requirements; the structural parameters of the initial laminate honeycomb sandwich structure are optimized, wherein the structural parameters include at least one of the number of laminate plies, the laminate ply direction, the laminate ply ratio, the laminate size, and the honeycomb sandwich thickness. at least one; using the laminate honeycomb sandwich structure after the structural parameters are optimized as the next initial laminate honeycomb sandwich structure, and based on the next initial laminate honeycomb sandwich structure, updating the natural grid finite element model of the aircraft body structure, and based on the updated natural grid finite element model of the aircraft body structure, continuing to judge whether the next initial laminate honeycomb sandwich structure meets the strength design requirements, stiffness design requirements, and weight design requirements, until the next initial laminate honeycomb sandwich structure meets the strength design requirements, stiffness design requirements, and weight design requirements at the same time, then determining the next initial laminate honeycomb sandwich structure as the frozen laminate honeycomb sandwich structure applied to the aircraft body structure.

[0108] Specifically, if the initial laminate honeycomb sandwich structure does not meet the design requirements, the initial laminate honeycomb sandwich structure is optimized in terms of the number of laminate layers, laminate layer direction, laminate layer ratio, laminate size, honeycomb sandwich thickness and other structural parameters to obtain an optimized laminate honeycomb sandwich structure, and the optimized laminate honeycomb sandwich structure is further determined by the method of steps 201-206 to determine whether the optimized laminate honeycomb sandwich structure meets the design requirements. If it does, the optimized laminate honeycomb sandwich structure is determined to be the laminate honeycomb sandwich structure ultimately used for the aircraft fuselage, otherwise the optimized laminate honeycomb sandwich structure is further optimized until the optimized laminate honeycomb sandwich structure simultaneously meets the strength design requirements, the stiffness design requirements and the weight design requirements. The three-dimensional finite element model of the aircraft can be used to accurately analyze the laminate honeycomb sandwich structure and predict its response characteristics under various working conditions, thereby improving the design accuracy. The finite element model supports rapid iterative calculations, and the design schemes can be optimized, evaluated and compared in a short time, thereby improving design efficiency and reducing costs. Based on the defined structural design variables and constraints, and taking into account strength, stiffness, manufacturing and weight factors, a multi-scale coupled structural size feasible design was carried out, achieving rapid selection and optimization of the size of the laminated honeycomb sandwich structure.

[0109] Furthermore, after determining the frozen laminate honeycomb sandwich structure that will be ultimately used on the aircraft, in order to ensure that the frozen laminate honeycomb sandwich structure meets the process requirements, it is also necessary to conduct process compliance inspections such as material information inspection, tooling inspection, stacking process inspection, curing process inspection, demolding process inspection, non-destructive testing inspection, CNC finishing inspection, and repair method inspection on the frozen laminate honeycomb sandwich structure, so as to ensure the flight stability and safety of the aircraft.

[0110] According to another design method of an aircraft laminate honeycomb sandwich structure provided by the present invention, compared with the prior art, in the embodiment of the present invention, by substituting the laminate honeycomb sandwich structure into the overall structure of the aircraft for stress state analysis during the process of optimizing the design of the laminate honeycomb sandwich structure, not only the stress state of the laminate honeycomb sandwich structure in the external environment alone can be analyzed, but also the structural stress state between the laminate honeycomb sandwich structure and other structures of the aircraft fuselage can be analyzed, so that the stress state analysis of the laminate honeycomb sandwich structure is more comprehensive, thereby improving the design accuracy of the laminate honeycomb sandwich structure. At the same time, by simulating the actual flight state of the aircraft to obtain the stress state of the laminate honeycomb sandwich structure, the complex and changeable service environment of the aircraft can be considered, thereby ensuring that the stress state obtained by the analysis is more in line with the actual aircraft conditions, thereby further improving the design accuracy of the laminate honeycomb sandwich structure. Furthermore, in the optimization design of the laminate honeycomb sandwich structure, the present invention comprehensively considers various failure modes related to the design, thereby improving the design accuracy of the laminate honeycomb sandwich structure.

[0111] Further, as Figure 1 In a specific implementation, an embodiment of the present invention provides a design device for a honeycomb sandwich structure of an aircraft laminate, such as Figure 6 As shown, the device includes: an acquisition unit 31, a first determination unit 32, an application unit 33, a second determination unit 34, and a judgment unit 35.

[0112] The acquisition unit 31 can be used to determine the initial laminate honeycomb sandwich structure of the laminate honeycomb sandwich to be designed, and to obtain the body property data of the aircraft body structure with the initial laminate honeycomb sandwich structure, and to construct a natural grid finite element model of the aircraft body structure based on the body property data, wherein the body property data includes the structural property data of the initial laminate honeycomb sandwich structure.

[0113] The first determination unit 32 can be used to determine the actual load, actual displacement constraint conditions, and actual flight environment conditions that the actual aircraft is borne during service, and to determine a simulated load that is the same as the actual load, a simulated displacement constraint condition that is the same as the actual displacement constraint condition, and a simulated flight environment condition that is the same as the actual flight environment condition.

[0114] The applying unit 33 can be used to apply the simulated load, set the simulated displacement constraint conditions, set the simulated flight environment conditions to the natural grid finite element model of the aircraft body structure, and determine the stress state data of the initial laminate honeycomb sandwich structure after applying the simulated load, setting the simulated displacement constraint conditions, and setting the simulated flight environment conditions.

[0115] The second determination unit 34 can be used to determine multiple failure modes of the initial laminate honeycomb sandwich structure based on the stress state data, and determine the structural design evaluation parameters of the initial laminate honeycomb sandwich structure under each failure mode based on the stress state data.

[0116] The judgment unit 35 can be used to judge whether the initial laminate honeycomb sandwich structure meets the design requirements based on the various structural design evaluation parameters and the structural property data. If the initial laminate honeycomb sandwich structure meets the design requirements, the initial laminate honeycomb sandwich structure is determined as a frozen laminate honeycomb sandwich structure applied to the aircraft fuselage structure. Otherwise, the structural parameters of the initial laminate honeycomb sandwich structure are optimized.

[0117] In a specific application scenario, in order to determine the initial laminate honeycomb sandwich structure, the acquisition unit 31 can be specifically used to determine the initial structural design requirements of the laminate of the laminate to be designed in the laminate honeycomb sandwich to be designed and the initial structural design requirements of the honeycomb sandwich to be designed, wherein the initial structural design requirements of the laminate include: at least two fiber directions are arranged, the ply in contact with the core and the surface ply need to be continuous, an anti-permeation film is laid, the layer is lost in the core area, the first layer is lost The distance from the honeycomb edge is greater than the first preset distance threshold, and the minimum distance between two removed layers is greater than the second preset distance threshold. The initial structural design requirements of the honeycomb sandwich include: the direction of the honeycomb core band is parallel to the main load direction, the radius of the honeycomb corner area is greater than the preset radius threshold, the honeycomb bevel angle is the preset angle threshold, the height of the honeycomb bevel edge is greater than the preset height threshold, a laminate sleeve is embedded in the honeycomb core, and adhesives and foam cores of preset materials are used on the inside and outside of the honeycomb core; based on the initial structural design requirements of the laminate and the initial structural design requirements of the honeycomb sandwich of the honeycomb sandwich to be designed, the initial laminate honeycomb sandwich structure is determined.

[0118] In a specific application scenario, in order to construct a natural grid finite element model of an aircraft fuselage structure, the acquisition unit 31 includes a determination module 311 and a construction module 312 .

[0119] The determination module 311 can be used to determine the coordinate system information, unit information, and grid division strategy information required for constructing the natural grid finite element model of the aircraft body structure, wherein the coordinate system information is to select the overall coordinate system of the aircraft body structure, define a spatial point on the aircraft symmetry plane directly in front of the aircraft nose as the origin O in the overall coordinate system, the x-axis in the overall coordinate system points to the left side of the aircraft along the heading, the x-axis is perpendicular to the aircraft fuselage along the heading, and the y-axis in the overall coordinate system passes through the center of the arc of the typical cross-section of the aircraft fuselage, and the y-axis points to the left side of the aircraft along the heading. On the opposite side of the heading, the z-axis follows the right-hand rule and passes through the origin O vertically upward. The meshing strategy information includes mesh nodes, unit numbers, unit attributes, and mesh sizes. The mesh nodes are selected at the intersections of the theoretical outer shape surface of the skin of the aircraft body structure, the theoretical outer shape plane of the beam, and the mid-surface of the long stringer, the mid-surface of the rib, and the mid-surface of the frame. The unit numbering adopts beam elements, rod elements, shell elements, supplemented by rigid units, constraint units, scalar units, and three-dimensional volume elements. The initial laminate honeycomb sandwich structure adopts shell element discretization, and the initial laminate honeycomb sandwich structure adopts a layering method to define unit attributes.

[0120] The construction module 312 can be used to construct the natural grid finite element model of the aircraft fuselage structure using preset finite element modeling software based on the fuselage structure size data, fuselage material property data, coordinate system information, unit information, and grid division strategy information.

[0121] In a specific application scenario, in order to apply actual flight conditions to the natural grid finite element model of the aircraft body structure, the applying unit 33 includes an applying module 331 and a setting module 332 .

[0122] The application module 331 can be used to apply the aerodynamic load and the inertial load to the rib, frame, and beam station surface skin nodes of the natural grid finite element model of the aircraft body structure respectively, and the concentrated load is applied to the corresponding structural unit through the rigid body unit.

[0123] The applying module 331 can also be used to select any three nodes of the aircraft in the natural grid finite element model of the aircraft body structure to apply node displacement constraints, apply the aircraft front landing gear displacement constraints in the x-axis and z-axis directions in the overall coordinate system at the aircraft front landing gear in the natural grid finite element model of the aircraft body structure, and apply the aircraft rear landing gear displacement constraints in the x-axis, y-axis and z-axis directions in the overall coordinate system at the two aircraft rear landing gears in the natural grid finite element model of the aircraft body structure.

[0124] The setting module 332 may be used to set the humidity condition and the temperature condition for the natural grid finite element model of the aircraft fuselage structure.

[0125] In a specific application scenario, in order to determine the structural design evaluation parameters under each of the failure modes, the second determination unit 34 can be specifically used to determine the honeycomb core crush evaluation parameters under the honeycomb core crush failure mode based on the stress in the thickness direction of the honeycomb core in the initial laminate honeycomb sandwich structure and the compressive strength of the honeycomb sandwich material; determine the honeycomb core shear tear evaluation parameters under the honeycomb core shear tear failure mode based on the longitudinal shear stress and transverse shear stress of the honeycomb sandwich in the initial laminate honeycomb sandwich structure, and the longitudinal shear strength and transverse shear strength of the honeycomb sandwich material; determine the honeycomb core shear tear evaluation parameters under the honeycomb core shear tear failure mode based on the x-axis stress in the x-axis direction and the y-axis stress in the y-axis direction of the laminates in the initial laminate honeycomb sandwich structure. The laminate failure evaluation parameters under the laminate failure failure mode are determined based on the y-axis stress, the x-axis shear stress corresponding to the x-axis and y-axis directions, the x-axis allowable stress in the x-axis direction of the prepreg corresponding to the laminate in the overall coordinate system, the y-axis allowable stress in the y-axis direction, and the xy-axis shear allowable stress corresponding to the x-axis and y-axis directions; the Poisson's ratio of the laminate in the initial laminate honeycomb sandwich structure, the x-axis compressive elastic modulus of the laminate along the x-axis direction, the y-axis compressive elastic modulus along the y-axis direction, the shear modulus of the laminate in the xy plane where the x-axis and y-axis are located, the laminate thickness of the laminate, and the honeycomb inner cell of the honeycomb sandwich structure in the initial laminate honeycomb sandwich structure. The lattice buckling evaluation parameters under the lattice buckling failure mode are determined based on the diameter of the cutting circle, the x-axis stress of the honeycomb wall panel unit of the honeycomb sandwich structure in the x-axis direction, the y-axis stress in the y-axis direction, the lattice buckling shear design allowable value of the honeycomb sandwich wall panel, the shear stress of the honeycomb wall panel unit, the first calculation parameter, and the second calculation parameter, wherein the lattice buckling evaluation parameters include: the lattice buckling evaluation parameters under the compression load, the lattice buckling evaluation parameters under the shear load, and the lattice buckling evaluation parameters under the combined action of the compression load and the shear load; based on the length of the laminate in the load direction, the thickness of the laminate, the bending stiffness coefficient of the laminate, the half-wave number in the load direction, and the angle perpendicular to the load direction of the laminate in the initial laminate honeycomb sandwich structure. The panel wrinkling evaluation parameters under the panel wrinkling failure mode are determined based on the width of the laminate in the loading direction, the panel centroid spacing of the laminate, the z-axis compressive elastic modulus of the laminate along the z-axis direction, the x-axis stress of the honeycomb wall panel unit of the honeycomb sandwich structure in the x-axis direction, and the y-axis stress in the y-axis direction; the shear wrinkle evaluation parameters under the shear wrinkle failure mode are determined based on the honeycomb thickness of the honeycomb sandwich structure in the initial laminate honeycomb sandwich structure, the laminate thickness of different laminates, the equivalent elastic modulus of different laminates, the shear modulus of the honeycomb sandwich in the xy plane where the x-axis and y-axis are located, the x-axis stress of the honeycomb wall panel unit of the honeycomb sandwich structure in the x-axis direction, and the y-axis stress in the y-axis direction;Based on the critical buckling load and the actual bearing load of the initial laminate honeycomb sandwich structure, the overall buckling evaluation parameters under the overall buckling failure mode are determined. ;

[0126] In a specific application scenario, in order to determine whether the initial laminate honeycomb sandwich structure meets the design requirements, the determination unit 35 includes a determination module 351 and a determination module 352 .

[0127] The judgment module 351 can be used to judge whether the initial laminate honeycomb sandwich structure meets the strength design requirements based on the honeycomb core crushing evaluation parameter, the honeycomb core shear and tearing evaluation parameter, the laminate failure evaluation parameter, the lattice buckling evaluation parameter, the panel wrinkling evaluation parameter, the shear wrinkle evaluation parameter, and the overall buckling evaluation parameter; judge whether the initial laminate honeycomb sandwich structure meets the stiffness design requirements based on the honeycomb core thickness, the honeycomb core shear modulus, the bending stiffness, and the shear stiffness; and judge whether the initial laminate honeycomb sandwich structure meets the weight design requirements based on the structure weight. The method for judging whether the initial laminate honeycomb sandwich structure meets the strength design requirements includes: judging whether the honeycomb core crushing evaluation parameter, the honeycomb core shear and tearing evaluation parameter, the laminate failure evaluation parameter, the lattice buckling evaluation parameter, the panel wrinkling evaluation parameter, the shear wrinkle evaluation parameter, and the overall buckling evaluation parameter are all greater than a preset parameter threshold; if they are all greater than the preset parameter threshold, then it is judged that the initial laminate honeycomb sandwich structure meets the strength design requirements.

[0128] The determination module 352 may be used to determine that the initial laminate honeycomb sandwich structure meets the design requirements if the initial laminate honeycomb sandwich structure meets the strength design requirement, the stiffness design requirement, and the weight design requirement at the same time.

[0129] In a specific application scenario, in order to perform structural optimization on the initial laminate honeycomb sandwich structure, the device further includes an optimization unit 36 ​​.

[0130] The optimization unit 36 ​​can be used to determine that the initial laminate honeycomb sandwich structure does not meet the design requirements if the initial laminate honeycomb sandwich structure does not meet any one of the strength design requirements, the stiffness design requirements, and the weight design requirements; optimize the structural parameters of the initial laminate honeycomb sandwich structure, wherein the structural parameters include at least one of the number of laminate plies, the laminate ply direction, the laminate ply ratio, the laminate size, and the honeycomb sandwich thickness; and use the laminate honeycomb sandwich structure after the structural parameter optimization as the next initial laminate honeycomb sandwich structure. structure, and based on the next initial laminate honeycomb sandwich structure, update the natural grid finite element model of the aircraft body structure, and based on the updated natural grid finite element model of the aircraft body structure, continue to judge whether the next initial laminate honeycomb sandwich structure meets the strength design requirements, stiffness design requirements, and weight design requirements, until the next initial laminate honeycomb sandwich structure satisfies the strength design requirements, stiffness design requirements, and weight design requirements at the same time, then the next initial laminate honeycomb sandwich structure is determined as the frozen laminate honeycomb sandwich structure applied to the aircraft body structure.

[0131] In a specific application scenario, in order to inspect the frozen laminate honeycomb sandwich structure, the device also includes an inspection unit 37 .

[0132] The inspection unit 37 can be used to perform at least one of material information inspection, tooling inspection, stacking process inspection, curing process inspection, demoulding process inspection, non-destructive testing inspection, CNC finishing inspection, and repair method inspection on the frozen laminate honeycomb sandwich structure.

[0133] It should be noted that for other corresponding descriptions of the functional modules involved in the design device for an aircraft laminate honeycomb sandwich structure provided by an embodiment of the present invention, reference can be made to Figure 1 The corresponding description of the method shown will not be repeated here.

[0134] Based on the above Figure 1The method shown, accordingly, an embodiment of the present invention further provides a computer-readable storage medium, on which a computer program is stored, and when the program is executed by a processor, the following steps are implemented: determining an initial laminate honeycomb sandwich structure of the laminate honeycomb sandwich to be designed, and obtaining body property data of the aircraft fuselage structure with the initial laminate honeycomb sandwich structure, and constructing a natural grid finite element model of the aircraft fuselage structure based on the body property data, wherein the body property data includes the structural property data of the initial laminate honeycomb sandwich structure; determining the actual load, actual displacement constraint conditions, and actual flight environment conditions borne by the actual aircraft during service, determining a simulated load with the same load as the actual load, a simulated displacement constraint condition with the same constraint as the actual displacement constraint condition, and a simulated flight environment condition with the same environment as the actual flight environment condition; applying a natural grid finite element model of the aircraft fuselage structure to the natural grid finite element model of the aircraft fuselage structure. The simulated load is applied, the simulated displacement constraint condition is set, the simulated flight environment condition is set, and the stress state data of the initial laminate honeycomb sandwich structure after applying the simulated load, setting the simulated displacement constraint condition, and setting the simulated flight environment condition is determined; based on the stress state data, multiple failure modes of the initial laminate honeycomb sandwich structure are determined, and based on the stress state data, the structural design evaluation parameters of the initial laminate honeycomb sandwich structure under each failure mode are determined; based on each of the structural design evaluation parameters and the structural property data, whether the initial laminate honeycomb sandwich structure meets the design requirements is judged, if the initial laminate honeycomb sandwich structure meets the design requirements, the initial laminate honeycomb sandwich structure is determined as a frozen laminate honeycomb sandwich structure applied to an aircraft fuselage structure, otherwise, the structural parameters of the initial laminate honeycomb sandwich structure are optimized.

[0135] Based on the above Figure 1 The method shown and Figure 6 The embodiment of the device shown in the figure, the embodiment of the present invention also provides a physical structure diagram of a computer device, such as Figure 7As shown, the computer device includes: a processor 41, a memory 42, and a computer program stored in the memory 42 and executable on the processor, wherein the memory 42 and the processor 41 are both arranged on a bus 43; and when the processor 41 executes the program, the following steps are implemented: determining an initial laminate honeycomb sandwich structure of the laminate honeycomb sandwich to be designed, and obtaining body property data of an aircraft fuselage structure with the initial laminate honeycomb sandwich structure, and constructing a natural grid finite element model of the aircraft fuselage structure based on the body property data, wherein the body property data includes structural property data of the initial laminate honeycomb sandwich structure; determining actual loads, actual displacement constraints, and actual flight environment conditions borne by the actual aircraft during service, determining simulated loads with the same load as the actual load, simulated displacement constraints with the same constraints as the actual displacement constraints, and simulated flight environment conditions with the same environment as the actual flight environment conditions; and The natural grid finite element model of the aircraft body structure applies the simulated load, sets the simulated displacement constraint condition, sets the simulated flight environment condition, and determines the stress state data of the initial laminate honeycomb sandwich structure after applying the simulated load, setting the simulated displacement constraint condition, and setting the simulated flight environment condition; based on the stress state data, determines a plurality of failure modes of the initial laminate honeycomb sandwich structure, and based on the stress state data, determines the structural design evaluation parameters of the initial laminate honeycomb sandwich structure under each of the failure modes; based on each of the structural design evaluation parameters and the structural property data, determines whether the initial laminate honeycomb sandwich structure meets the design requirements; if the initial laminate honeycomb sandwich structure meets the design requirements, determines the initial laminate honeycomb sandwich structure as a frozen laminate honeycomb sandwich structure applied to the aircraft body structure, otherwise, optimizes the structural parameters of the initial laminate honeycomb sandwich structure.

[0136] Through the technical solution of the present invention, the present invention constructs a natural grid finite element model of an aircraft fuselage structure with an initial laminate honeycomb sandwich structure, and applies the same simulated loads, simulated displacement constraints and simulated environmental conditions as those borne by the aircraft during actual flight service to the natural grid finite element model, and based on the above-mentioned simulated loads, simulated displacement constraints and simulated environmental conditions, simulates the aircraft service process, and based on the simulation results, determines the stress state of the initial laminate honeycomb sandwich structure after applying the simulated loads, simulated displacement constraints and simulated flight environmental conditions, and determines the failure mode according to the stress state, and calculates the evaluation parameters under each failure mode, and finally optimizes the initial laminate honeycomb sandwich structure according to the evaluation parameters under each failure mode and the attribute data of the initial laminate honeycomb sandwich structure, thereby, in the process of optimizing the design of the laminate honeycomb sandwich structure, Substituting the laminate honeycomb sandwich structure into the overall structure of the aircraft for stress state analysis can not only analyze the stress state of the laminate honeycomb sandwich structure in the external environment alone, but also analyze the structural stress state between the laminate honeycomb sandwich structure and other structures of the aircraft fuselage, so that the stress state analysis of the laminate honeycomb sandwich structure is more comprehensive, thereby improving the design accuracy of the laminate honeycomb sandwich structure. At the same time, by simulating the actual flight state of the aircraft to obtain the stress state of the laminate honeycomb sandwich structure, the complex and changeable service environment of the aircraft can be considered, thereby ensuring that the stress state obtained by analysis is more in line with the actual aircraft conditions, thereby further improving the design accuracy of the laminate honeycomb sandwich structure. Furthermore, the present invention comprehensively considers various failure modes related to the design in the optimal design of the laminate honeycomb sandwich structure, thereby improving the design accuracy of the laminate honeycomb sandwich structure.

[0137] Obviously, those skilled in the art should understand that the above modules or steps of the present invention can be implemented by a general computing device, they can be concentrated on a single computing device, or distributed on a network composed of multiple computing devices, and optionally, they can be implemented by a program code executable by a computing device, so that they can be stored in a storage device and executed by the computing device, and in some cases, the steps shown or described can be executed in a different order than here, or they can be made into individual integrated circuit modules, or multiple modules or steps therein can be made into a single integrated circuit module for implementation. Thus, the present invention is not limited to any specific combination of hardware and software.

[0138] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. For those skilled in the art, the present invention may have various modifications and variations. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

Claims

1. A design method for an aircraft laminate honeycomb sandwich structure, characterized in that: include: Determine the initial laminate honeycomb sandwich structure of the laminate honeycomb sandwich to be designed, obtain the body property data of the aircraft body structure with the initial laminate honeycomb sandwich structure, and construct a natural grid finite element model of the aircraft body structure based on the body property data, wherein the body property data includes the structural property data of the initial laminate honeycomb sandwich structure, the body structure size data of the aircraft body structure with the initial laminate honeycomb sandwich structure, and the body material property data. The method for constructing the natural grid finite element model of the aircraft body structure includes: determining the coordinate system information, unit information, and mesh division strategy information required for constructing the natural grid finite element model of the aircraft body structure; based on the body structure size data, body material property data, coordinate system information, unit information, and mesh division strategy information, construct the natural grid finite element model of the aircraft body structure using preset finite element modeling software; Determine the actual load, actual displacement constraint, and actual flight environment conditions borne by the actual aircraft during service, and determine a simulated load with the same load as the actual load, a simulated displacement constraint with the same constraint as the actual displacement constraint, and a simulated flight environment condition with the same environment as the actual flight environment condition; Applying the simulated load, setting the simulated displacement constraint condition, and setting the simulated flight environment condition to the natural grid finite element model of the aircraft body structure, and determining the stress state data of the initial laminate honeycomb sandwich structure after applying the simulated load, setting the simulated displacement constraint condition, and setting the simulated flight environment condition; Based on the stress state data, determining multiple failure modes of the initial laminate honeycomb sandwich structure, and based on the stress state data, determining structural design evaluation parameters of the initial laminate honeycomb sandwich structure under each of the failure modes; Based on the structural design evaluation parameters and the structural property data, it is determined whether the initial laminate honeycomb sandwich structure meets the design requirements. If the initial laminate honeycomb sandwich structure meets the design requirements, the initial laminate honeycomb sandwich structure is determined as a frozen laminate honeycomb sandwich structure applied to an aircraft fuselage structure. Otherwise, the structural parameters of the initial laminate honeycomb sandwich structure are optimized.

2. The method according to claim 1, characterized in that The step of determining an initial laminate honeycomb sandwich structure of the laminate honeycomb sandwich to be designed comprises: Determine the initial structural design requirements of the laminate of the laminate to be designed and the initial structural design requirements of the honeycomb sandwich of the honeycomb sandwich to be designed, wherein the initial structural design requirements of the laminate include: at least two fiber directions are arranged, the ply in contact with the core and the surface ply need to be continuous, an anti-permeation film is laid, the layer is dropped in the core area, the distance between the first dropped layer and the honeycomb edge is greater than a first preset distance threshold, and the minimum distance between two dropped layers is greater than a second preset distance threshold; the initial structural design requirements of the honeycomb sandwich include: the direction of the honeycomb core band is parallel to the main load direction, the radius of the honeycomb corner area is greater than a preset radius threshold, the honeycomb bevel angle is a preset angle threshold, the height of the honeycomb bevel edge is greater than a preset height threshold, a laminate liner is embedded in the honeycomb core, and adhesives and foam cores of preset materials are used inside and outside the honeycomb core; Based on the initial structural design requirements of the laminate and the initial structural design requirements of the honeycomb sandwich of the honeycomb sandwich to be designed, the initial laminate honeycomb sandwich structure is determined.

3. The method according to claim 1, characterized in that: The coordinate system information is to select the overall coordinate system of the aircraft body structure, define a spatial point on the aircraft symmetry plane directly in front of the aircraft nose as the origin O in the overall coordinate system, the x-axis in the overall coordinate system points to the left side of the aircraft along the heading, the x-axis is perpendicular to the aircraft fuselage along the heading, the y-axis in the overall coordinate system passes through the center of the arc of the typical cross-section of the aircraft fuselage, the y-axis points to the opposite side of the aircraft along the heading, and the z-axis follows the right-hand rule and passes through the origin O vertically upward. The grid division strategy information includes grid nodes, unit numbers, unit attributes, and grid sizes. The grid nodes are selected at the intersection of the theoretical skin shape surface, the theoretical beam shape plane, and the mid-surface of the long stringer and the mid-surface of the rib and the mid-surface of the frame of the aircraft body structure. The unit numbering adopts beam elements, rod elements, shell elements, supplemented by rigid units, constraint units, scalar units, and three-dimensional volume elements. The initial laminate honeycomb sandwich structure adopts shell element discretization, and the initial laminate honeycomb sandwich structure adopts a layering method to define unit attributes.

4. The method according to claim 3, characterized in that The simulated loads include: aerodynamic loads, inertial loads, concentrated loads; the simulated displacement constraints include: node displacement constraints, aircraft front landing gear displacement constraints, aircraft rear landing gear displacement constraints; the simulated flight environment conditions include humidity conditions and temperature conditions; The applying the simulated load to the natural grid finite element model of the aircraft body structure, setting the simulated displacement constraint condition, and setting the simulated flight environment condition comprises: The aerodynamic load and the inertial load are applied to the rib, frame, and beam station surface skin nodes of the natural grid finite element model of the aircraft body structure respectively, and the concentrated load is applied to the corresponding structural unit through the rigid body unit; Select any three nodes of the aircraft in the natural grid finite element model of the aircraft body structure to apply node displacement constraints, apply the displacement constraints of the front landing gear of the aircraft in the x-axis direction and the z-axis direction in the overall coordinate system to the front landing gear of the aircraft in the natural grid finite element model of the aircraft body structure, and apply the displacement constraints of the rear landing gear of the aircraft in the x-axis, y-axis and z-axis directions in the overall coordinate system to the two rear landing gears of the aircraft in the natural grid finite element model of the aircraft body structure; The humidity condition and the temperature condition are set for the aircraft fuselage structure natural grid finite element model.

5. The method according to claim 1, characterized in that The multiple failure modes include honeycomb core crushing failure mode, honeycomb core shear tearing failure mode, laminate damage failure mode, lattice buckling failure mode, panel wrinkling failure mode, shear wrinkle failure mode, and overall buckling failure mode; Determining the structural design evaluation parameters of the initial laminate honeycomb sandwich structure under each failure mode based on the stress state data includes: Determining a honeycomb core crushing evaluation parameter under the honeycomb core crushing failure mode based on the stress in the honeycomb core thickness direction of the initial laminate honeycomb sandwich structure and the compressive strength of the honeycomb core material; Determine the honeycomb core shear tear evaluation parameters under the honeycomb core shear tear failure mode based on the longitudinal shear stress and transverse shear stress of the honeycomb sandwich core in the initial laminate honeycomb sandwich structure and the longitudinal shear strength and transverse shear strength of the honeycomb sandwich material; Based on the x-axis stress in the x-axis direction, the y-axis stress in the y-axis direction, and the xy-axis shear stress corresponding to both the x-axis direction and the y-axis direction of the laminate in the initial laminate honeycomb sandwich structure in the overall coordinate system, the x-axis allowable stress in the x-axis direction, the y-axis allowable stress in the y-axis direction, and the xy-axis shear allowable stress corresponding to both the x-axis direction and the y-axis direction of the prepreg corresponding to the laminate in the overall coordinate system, determine the laminate failure evaluation parameters under the laminate failure mode; Based on the Poisson's ratio of the laminate in the initial laminate honeycomb sandwich structure, the x-axis compressive elastic modulus of the laminate along the x-axis direction, the y-axis compressive elastic modulus along the y-axis direction, the shear modulus of the laminate in the xy plane where the x-axis and the y-axis are located, the laminate thickness of the laminate, the honeycomb inscribed circle diameter of the honeycomb sandwich structure in the initial laminate honeycomb sandwich structure, the x-axis stress of the honeycomb wall panel unit in the x-axis direction, the y-axis stress in the y-axis direction, the lattice buckling shear design allowable value of the honeycomb sandwich wall panel, the shear stress of the honeycomb wall panel unit, the first calculation parameter, and the second calculation parameter, determine the lattice buckling evaluation parameters under the lattice buckling failure mode, wherein the lattice buckling evaluation parameters include: the lattice buckling evaluation parameters under the compression load, the lattice buckling evaluation parameters under the shear load, and the lattice buckling evaluation parameters under the combined action of the compression load and the shear load; Based on the length of the laminate in the initial laminate honeycomb sandwich structure in the loading direction, the laminate thickness, the laminate bending stiffness coefficient, the half-wave number in the loading direction, the laminate width perpendicular to the loading direction, the panel centroid spacing of the laminate, the z-axis compressive elastic modulus of the laminate along the z-axis direction, the x-axis stress of the honeycomb wall panel unit of the honeycomb sandwich structure in the x-axis direction, and the y-axis stress in the y-axis direction, determine the panel wrinkling evaluation parameters under the panel wrinkling failure mode; Determine the shear wrinkle evaluation parameters under the shear wrinkle failure mode based on the honeycomb thickness of the honeycomb sandwich structure in the initial laminate honeycomb sandwich structure, the laminate thickness of different laminates, the equivalent elastic modulus of different laminates, the shear modulus of the honeycomb sandwich in the xy plane where the x-axis and the y-axis are located, the x-axis stress of the honeycomb wall panel unit of the honeycomb sandwich structure in the x-axis direction, and the y-axis stress in the y-axis direction; Based on the critical buckling load and the actual bearing load of the initial laminate honeycomb sandwich structure, the overall buckling evaluation parameters under the overall buckling failure mode are determined.

6. The method according to claim 5, characterized in that The structural attribute data include honeycomb core thickness, honeycomb core shear modulus, bending stiffness, shear stiffness, and structural weight corresponding to the initial laminate honeycomb sandwich structure; The step of judging whether the initial laminate honeycomb sandwich structure meets the design requirements based on the structural design evaluation parameters and the structural property data includes: Based on the honeycomb core crushing evaluation parameter, the honeycomb core shear tearing evaluation parameter, the laminate damage evaluation parameter, the lattice buckling evaluation parameter, the panel wrinkling evaluation parameter, the shear wrinkle evaluation parameter, and the overall buckling evaluation parameter, it is judged whether the initial laminate honeycomb sandwich structure meets the strength design requirement; based on the honeycomb core thickness, the honeycomb core shear modulus, the bending stiffness, and the shear stiffness, it is judged whether the initial laminate honeycomb sandwich structure meets the stiffness design requirement; based on the structure weight, it is judged whether the initial laminate honeycomb sandwich structure meets the weight design requirement, wherein the method for judging whether the initial laminate honeycomb sandwich structure meets the strength design requirement includes: judging whether the honeycomb core crushing evaluation parameter, the honeycomb core shear tearing evaluation parameter, the laminate damage evaluation parameter, the lattice buckling evaluation parameter, the panel wrinkling evaluation parameter, the shear wrinkle evaluation parameter, and the overall buckling evaluation parameter are all greater than a preset parameter threshold; if they are all greater than the preset parameter threshold, it is judged that the initial laminate honeycomb sandwich structure meets the strength design requirement; If the initial laminate honeycomb sandwich structure satisfies the strength design requirement, the stiffness design requirement, and the weight design requirement at the same time, it is determined that the initial laminate honeycomb sandwich structure meets the design requirements; After judging whether the initial laminate honeycomb sandwich structure meets the design requirements based on the structural design evaluation parameters and the structural attribute data, the method further includes: If the initial laminate honeycomb sandwich structure does not meet any one of the strength design requirement, the stiffness design requirement, and the weight design requirement, then it is determined that the initial laminate honeycomb sandwich structure does not meet the design requirement; Optimizing the structural parameters of the initial laminate honeycomb sandwich structure, wherein the structural parameters include at least one of the number of laminate plies, the direction of laminate plies, the ratio of laminate plies, the size of laminates, and the thickness of the honeycomb sandwich structure; The laminate honeycomb sandwich structure after the structural parameter optimization is used as the next initial laminate honeycomb sandwich structure, and based on the next initial laminate honeycomb sandwich structure, the natural grid finite element model of the aircraft body structure is updated, and based on the updated natural grid finite element model of the aircraft body structure, the next initial laminate honeycomb sandwich structure is continuously judged whether it meets the strength design requirements, stiffness design requirements, and weight design requirements, until the next initial laminate honeycomb sandwich structure satisfies the strength design requirements, stiffness design requirements, and weight design requirements at the same time, then the next initial laminate honeycomb sandwich structure is determined as the frozen laminate honeycomb sandwich structure applied to the aircraft body structure.

7. The method according to claim 1, characterized in that After determining the initial laminate honeycomb sandwich structure as a frozen laminate honeycomb sandwich structure applied to an aircraft fuselage structure, the method further comprises: The frozen laminate honeycomb sandwich structure is subjected to at least one of material information inspection, tooling inspection, stacking process inspection, curing process inspection, demoulding process inspection, non-destructive testing inspection, CNC finishing inspection, and repair method inspection.

8. A design device for an aircraft laminate honeycomb sandwich structure, characterized in that: include: An acquisition unit is used to determine an initial laminate honeycomb sandwich structure of a laminate honeycomb sandwich to be designed, and to acquire body attribute data of an aircraft body structure with the initial laminate honeycomb sandwich structure, and to construct a natural mesh finite element model of the aircraft body structure based on the body attribute data, wherein the body attribute data includes structural attribute data of the initial laminate honeycomb sandwich structure, body structure dimension data of the aircraft body structure with the initial laminate honeycomb sandwich structure, and body material attribute data. The method for constructing the natural mesh finite element model of the aircraft body structure includes: determining coordinate system information, unit information, and mesh division strategy information required for constructing the natural mesh finite element model of the aircraft body structure; and constructing the natural mesh finite element model of the aircraft body structure using a preset finite element modeling software based on the body structure dimension data, body material attribute data, coordinate system information, unit information, and mesh division strategy information. The first determination unit is used to determine the actual load, actual displacement constraint condition, and actual flight environment condition borne by the actual aircraft during service, and determine a simulated load with the same load as the actual load, a simulated displacement constraint condition with the same constraint as the actual displacement constraint condition, and a simulated flight environment condition with the same environment as the actual flight environment condition; an applying unit, used for applying the simulated load, setting the simulated displacement constraint condition, and setting the simulated flight environment condition to the natural grid finite element model of the aircraft body structure, and determining stress state data of the initial laminate honeycomb sandwich structure after applying the simulated load, setting the simulated displacement constraint condition, and setting the simulated flight environment condition; A second determination unit is used to determine multiple failure modes of the initial laminate honeycomb sandwich structure based on the stress state data, and determine a structural design evaluation parameter of the initial laminate honeycomb sandwich structure under each of the failure modes based on the stress state data; A judgment unit is used to judge whether the initial laminate honeycomb sandwich structure meets the design requirements based on each of the structural design evaluation parameters and the structural property data; if the initial laminate honeycomb sandwich structure meets the design requirements, the initial laminate honeycomb sandwich structure is determined as a frozen laminate honeycomb sandwich structure applied to an aircraft fuselage structure; otherwise, the structural parameters of the initial laminate honeycomb sandwich structure are optimized.

9. A computer-readable storage medium having a computer program stored thereon, characterized in that: When the computer program is executed by a processor, the steps of the method according to any one of claims 1 to 7 are implemented.

10. A computer device comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that: When the computer program is executed by a processor, the steps of the method according to any one of claims 1 to 7 are implemented.

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