Design method of unmanned aerial vehicle main load-bearing structure and related device

Abstract

This invention belongs to the field of UAV structural design technology, and discloses a design method and related apparatus for the main load-bearing structure of a UAV. The design method first obtains a wireframe skeleton model of the main load-bearing structure of the UAV, constructs an initial finite element model with single material properties and set initial cross-sectional dimensions, and iteratively optimizes the initial cross-sectional dimensions to obtain an optimized finite element model. Then, the materials of each functional structural region are restored to the target material, and a multi-material finite element model is constructed. Finally, based on the multi-material finite element model, the design result is obtained after multi-load condition analysis and adjustment. The technical solution disclosed in this invention reduces the complexity of multi-material parameter coupling and improves the iterative efficiency and lightweight reliability of the UAV main load-bearing structure design through an integrated hierarchical optimization strategy of material equivalence, size optimization, material restoration, and strength verification.

Description

Technical Field

[0001] This invention belongs to the field of unmanned aerial vehicle (UAV) structural design technology, and specifically relates to a design method and related device for the main load-bearing structure of an UAV. Background Technology

[0002] In recent years, against the backdrop of the deepening global technological revolution and industrial transformation, the low-altitude economy, as an important direction of the national strategic emerging industries, is ushering in unprecedented development opportunities. Unmanned aerial vehicles (UAVs), as an important carrier and key driving force in the low-altitude economic system, have become a focus of common attention in academia, industry, and policy research.

[0003] Currently, the traditional design method for the main load-bearing structure of UAVs typically employs a variety of materials for zoned design based on the different performance requirements of different functional areas. For example, the front fuselage uses glass fiber composite materials to balance wave transmission and lightweighting, the middle and rear fuselages use carbon fiber composite materials to achieve high specific strength and stiffness, the internal frame beams use high-strength aluminum alloys to ensure connection and support reliability, the heated areas near the engine use titanium alloys for high-temperature resistance, and alloy steel is used in locally high-load-bearing parts to ensure load-bearing safety. In the above-mentioned zoned design process, a finite element model needs to be established and strength, stiffness, and stability analyses need to be performed for various load conditions of the UAV (such as ground parking, taxiing, take-off and landing, and aerial maneuvering). Through iterative evaluation based on failure criteria (such as maximum stress, strain, or buckling critical load), parameters such as the cross-sectional dimensions, composite material lay-up angles, and thicknesses of each area are manually adjusted to pursue a balance between structural lightweighting and safety.

[0004] Traditional design methods for the main load-bearing structure of unmanned aerial vehicles (UAVs) still have the following problems, including: (1) Faced with hundreds of working condition combinations including aerodynamics, inertia, impact and thermal load, relying on manual experience to adjust and iterate parameters is inefficient and prone to getting trapped in local optima. (2) The coupling between different material properties, geometric dimensions and ply design is complex, and it is difficult to achieve system-level convergence in multiple iterations, which often leads to a lengthy design process and large fluctuations in results; (3) The application of Finite Element Analysis (FEA) technology generally suffers from the problem of “analysis lag”, that is: the relevant analysis functions are only used for performance verification and validation after the design scheme is finalized, and cannot be effectively intervened and positively guided in the conceptual design, scheme selection and structural iteration optimization stages; as the configuration of UAV structures becomes increasingly complex, the requirements for lightweighting continue to increase, and new material systems such as composite materials and smart materials are widely used, the above-mentioned “analysis lag” problem will further lead to problems such as high simulation calculation costs and long feedback cycles in the design iteration process, which seriously restricts the iteration efficiency of R&D design.

[0005] In summary, existing traditional design methods lack an integrated, hierarchical optimization strategy from materials to structure, resulting in problems such as poor design convergence and difficulty in guaranteeing reliability, which restricts further improvements in lightweight, high-performance and rapid development of UAV main load-bearing structures. Summary of the Invention

[0006] The purpose of this invention is to provide a design method and related apparatus for the main load-bearing structure of an unmanned aerial vehicle (UAV) to solve one or more of the aforementioned technical problems. The technical solution disclosed in this invention can be applied to the UAV structural design process, specifically providing an integrated, hierarchical optimization strategy encompassing "material equivalence, size optimization, material reduction, and strength verification." This strategy can effectively save time in defining structural cross-sectional dimensions and significantly improve the iterative efficiency of UAV structural design.

[0007] To achieve the above objectives, the present invention adopts the following technical solution: In a first aspect, the present invention provides a design method for the main load-bearing structure of an unmanned aerial vehicle (UAV), comprising the following steps: Obtain the wireframe skeleton model of the main load-bearing structure of the UAV; Based on the wireframe skeleton model, an initial finite element model of the main load-bearing structure of the UAV is constructed; wherein, in the initial finite element model, a single material property is assigned according to the material with the largest amount used in the UAV design, and the initial cross-sectional dimensions are set; Based on the selected multiple load conditions, the initial finite element model is loaded and subjected to finite element analysis to obtain stress-displacement analysis results. Based on the stress-displacement analysis results, the initial cross-sectional dimensions are iteratively adjusted and optimized so that the optimized cross-sectional dimensions meet all load conditions, and the optimized finite element model is obtained. Based on the optimized finite element model, the materials of each functional structural region in the main load-bearing structure of the UAV are restored to the target materials to obtain a multi-material finite element model. Based on the selected multiple load conditions, finite element analysis is performed on the multi-material finite element model to obtain stress-displacement analysis results. Based on the stress-displacement analysis results, adjustments are made, and the adjusted multi-material finite element model is used as the design result of the main load-bearing structure of the UAV.

[0008] A further improvement to the technical solution of this invention lies in that the step of obtaining the wireframe skeleton model of the main load-bearing structure of the UAV specifically includes: Obtain the design requirements for the main load-bearing structure of the UAV; wherein, the design requirements include: overall load conditions and airborne system layout; Based on the design requirements, design a truss structure model or a semi-rigid shell structure model to obtain the main load-bearing structure model of the UAV. Based on the main load-bearing structure model of the UAV, wireframe modeling is performed to obtain the wireframe skeleton model of the main load-bearing structure of the UAV. In the process of wireframe modeling, according to the overall load conditions and the layout of the airborne system, the rods or flanges that transmit torque and axial force are abstracted into line models, the surfaces that transmit in-plane shear loads or out-of-plane loads are abstracted into surface models, and the large mass blocks of the airborne system are abstracted into mass points. The connection relationship between the line models, surface models and mass points is established.

[0009] A further improvement of the technical solution of the present invention lies in that, in the process of designing a truss structure model or a semi-rigid shell structure model based on the design requirements to obtain the main load-bearing structure model of the UAV, the design principles of shortest structural force transmission path, lightweight structure and maintainability of airborne equipment are adopted when designing the truss structure model or the semi-rigid shell structure model.

[0010] A further improvement of the technical solution of the present invention is that, in the process of constructing the initial finite element model of the main load-bearing structure of the UAV based on the wireframe skeleton model, the main load-bearing components that bear the longitudinal and lateral load transfer functions in the wireframe skeleton model are geometrically modeled using beam elements; and the intermediate area formed by connecting and closing the main load-bearing components is geometrically modeled using shell elements.

[0011] A further improvement to the technical solution of this invention lies in the iterative adjustment and optimization of the initial cross-sectional dimensions based on the stress-displacement analysis results, so that the optimized cross-sectional dimensions meet all load conditions. The principles for iteratively adjusting and optimizing the initial cross-sectional dimensions during the process of obtaining the optimized finite element model include: The force transmission structure is statically determinate. When the stress level in a local area of ​​the component is higher than that in other areas and the difference is greater than or equal to the first preset difference threshold, or when the stress level in a local area of ​​the component is greater than or equal to the allowable value of the material design, the thickness or outer diameter of the component is increased until the local stress level of the component meets the requirements. When the stress level in a local area of ​​the component is lower than that in other areas and the difference is greater than or equal to the second preset difference threshold, the thickness or outer diameter of the component is reduced until the local stress level of the component meets the requirements. The force transmission structure is statically indeterminate. When the stress level in a local area of ​​the component is higher than that in other areas and the difference is greater than or equal to the third preset difference threshold, the component thickness or outer diameter is reduced until the local stress level of the component meets the requirements. When the stress level in a local area of ​​the component is lower than that in other areas and the difference is greater than or equal to the fourth preset difference threshold, the component thickness or outer diameter is increased until the local stress level of the component meets the requirements. Among them, the components meet the functional requirements of the corresponding functional area at the position of maximum displacement, and there is no interference between the functional components.

[0012] A further improvement to the technical solution of this invention lies in the step of restoring the materials of each functional structural region in the main load-bearing structure of the UAV to the target material based on the optimized finite element model, thereby obtaining a multi-material finite element model, including: Based on the division of each functional structural region, the property parameters of the target material are called and assigned to the corresponding unit set; Based on the principle of equivalent stiffness of structure, the cross-sectional dimensions of the element set are replaced by equivalent modeling to obtain a multi-material finite element model.

[0013] A further improvement to the technical solution of this invention lies in the following steps: Based on selected multiple load conditions, finite element analysis is performed on the multi-material finite element model to obtain stress-displacement analysis results. Adjustments are then made based on these results, and the adjusted multi-material finite element model is used as the design result for the main load-bearing structure of the UAV. After adjustments based on the stress-displacement analysis results, the following verification and adjustment steps are also included: The pressure-bearing components were checked and adjusted according to the instability failure criteria; The heated components were checked and adjusted according to the thermal stress failure criteria; Fatigue-dependent components should be checked and adjusted according to fatigue failure criteria.

[0014] In a second aspect, the present invention provides a design system for the main load-bearing structure of an unmanned aerial vehicle (UAV), comprising: The wireframe skeleton model acquisition module is used to acquire the wireframe skeleton model of the main load-bearing structure of the UAV. The initial finite element model construction module is used to construct an initial finite element model of the main load-bearing structure of the UAV based on the wireframe skeleton model; wherein, in the initial finite element model, a single material property is assigned according to the material with the largest amount used in the UAV design, and the initial cross-sectional dimensions are set; The optimized finite element model acquisition module is used to load and perform finite element analysis on the initial finite element model based on selected multiple load conditions, obtain stress-displacement analysis results, and iteratively adjust and optimize the initial cross-sectional dimensions based on the stress-displacement analysis results so that the optimized cross-sectional dimensions meet all load conditions, thereby obtaining the optimized finite element model. The design result acquisition module is used to restore the materials of each functional structural region in the main load-bearing structure of the UAV to the target materials based on the optimized finite element model, and obtain a multi-material finite element model; based on the selected multiple load conditions, perform finite element analysis on the multi-material finite element model to obtain stress-displacement analysis results; adjust based on the stress-displacement analysis results; and use the adjusted multi-material finite element model as the design result of the main load-bearing structure of the UAV.

[0015] In a third aspect, the present invention provides an electronic device including a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor executes the computer program to implement the design method for the main load-bearing structure of an unmanned aerial vehicle as described in any one of the first aspects of the present invention.

[0016] In a fourth aspect, the present invention provides a non-transitory computer-readable storage medium having a computer program stored thereon, wherein the computer program, when executed by a processor, implements the design method for the main load-bearing structure of an unmanned aerial vehicle as described in any one of the first aspects of the present invention.

[0017] Compared with the prior art, the present invention has the following beneficial effects: This invention specifically discloses a design method for the main load-bearing structure of an unmanned aerial vehicle (UAV). The method first establishes a parametric finite element model using the same type of material in the preliminary definition stage of the structural cross-sectional dimensions. Then, based on the selected actual operating conditions, the stress levels at different structural locations under various operating conditions are calculated, and the cross-sectional dimensions are iteratively optimized (interpretably based on preset failure criteria such as strength and stability). Finally, the iteratively optimized cross-sectional parameters are mapped to different materials selected in the actual design, and comprehensive checks and adjustments are made regarding strength, stiffness, and failure modes. The final design result is obtained through these checks and adjustments. The novel method disclosed in this invention, through an integrated and hierarchical optimization strategy of "material equivalence, size optimization, material restoration, and strength verification," can significantly reduce the design coupling complexity introduced by the correlation of multiple material parameters, and improve the iterative efficiency and reliability of lightweight structural design. Specifically, this invention performs mechanical equivalence on selected materials and multiple metal structures in the initial optimization stage, performs systematic and efficient multi-condition topology and morphology optimization at the size level, restores the real material properties and refines the layup parameters, and finally verifies its load-bearing capacity through high-precision strength verification. This phased and hierarchical integrated optimization approach not only significantly improves the design convergence efficiency of complex multi-material structures under multiple constraints, but also effectively ensures the engineering feasibility of the final designed structure. Attached Figure Description

[0018] To more clearly illustrate the technical solutions in this invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of this invention. For those skilled in the art, other drawings can be obtained from these drawings without creative effort.

[0019] Figure 1 This is a flowchart illustrating a design method for the main load-bearing structure of an unmanned aerial vehicle (UAV) in an embodiment of the present invention.

[0020] Figure 2This is a schematic diagram of a design system for the main load-bearing structure of an unmanned aerial vehicle (UAV) in an embodiment of the present invention. Detailed Implementation

[0021] To make the objectives, technical solutions, and advantages of the present invention clearer, the technical solutions of the present invention will be clearly and completely described below with reference to the accompanying drawings of the embodiments of the present invention; obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments.

[0022] Based on the technical solutions disclosed in the embodiments of this invention, all other embodiments obtained by those skilled in the art without inventive effort are within the scope of protection of this invention. Furthermore, the terms "comprising" and "having," and any variations thereof, are intended to cover non-exclusive inclusion. For example, a process, method, system, product, or device that includes a series of steps or units is not necessarily limited to those steps or units explicitly listed, but may include other steps or units not explicitly listed or inherent to these processes, methods, products, or devices.

[0023] In the design of the main load-bearing structure of UAVs (such as frame beams), the use of different materials can lead to significant differences in the dimensional parameters and strength failure criteria of the cross-sections at various stations. This difference makes the cross-sectional dimension optimization process extremely complex when the structure is subjected to hundreds of load conditions. Manual iterative calculations are inefficient and severely restrict the convergence speed and engineering feasibility of lightweight designs. To address the technical challenges of increased design variables and complex parameter coupling relationships caused by material differences in UAV structural design, this invention discloses a design method for the main load-bearing structure of UAVs. Specifically, it is a graded optimization method for cross-sectional dimensions based on an equivalent material model, which can effectively save time in defining structural cross-sectional dimensions and significantly improve the iterative efficiency of UAV structural design.

[0024] Please see Figure 1 The present invention provides a design method for the main load-bearing structure of an unmanned aerial vehicle (UAV), comprising the following steps: Step 1: Obtain the wireframe skeleton model of the main load-bearing structure of the UAV.

[0025] In a specific exemplary technical solution, the drone is a rotary-wing drone. The specific process for obtaining the wireframe skeleton model of the drone's main load-bearing structure is as follows: Step 1.1: Obtain the design requirements for the main load-bearing structure of the UAV. For illustrative purposes, the design requirements specifically include: overall load conditions and onboard system layout. The overall load conditions may include: ground parking, ground gust, ground landing, ground mooring, ground hoisting, hovering, symmetrical maneuvers, discrete gusts, fixed-point turns, and large-scale yaw maneuvers. The onboard system layout includes: the center of mass, mass, and position information of large mass blocks such as the power unit, cooling system, fuel tank system, onboard power supply, flight control hardware, electrical control box, generator, antenna module, and data link.

[0026] In a further exemplary technical solution, the ground landing condition is explained as follows: In the UAV ground landing condition, the vertical impact load acts on the landing gear touchdown point, with the direction perpendicular to the contact surface. Its magnitude corresponds to the maximum vertical impact force at the moment of landing, forming a vertical load condition on the landing gear and fuselage; when the UAV touches down with forward flight speed, the ground friction force acts on the wheels or skids, with the direction opposite to the taxiing direction, forming a longitudinal shear load on the landing gear and fuselage connection structure; in asymmetrical landing or with lateral velocity, the lateral friction force of the ground acting on the landing gear constitutes a lateral bearing condition of the structure.

[0027] In a further exemplary technical solution, the hovering condition is explained as follows: In the vertical direction, the total thrust generated by the rotor system acts perpendicularly on the rotor shaft center and is transmitted to the main fuselage structure through the main reducer support. This thrust and the weight of the UAV form a pair of equal and opposite balanced force systems, forming the main vertical load condition of the fuselage. In terms of torque, the aerodynamic torque generated by the rotor rotation acts directly on the rotor shaft system and the main reducer mounting interface. This torque constitutes the torsional load condition on the fuselage. The lateral thrust generated by the tail rotor acts on the tail rotor axis center and is transmitted to the fuselage through the tail boom structure. The yaw control torque formed therefrom and the aforementioned rotor torque form a pair of opposite balanced torques, forming the yaw direction torque balance load condition of the fuselage. Forces and torques in other directions can be ignored.

[0028] Step 1.2: Based on the design requirements obtained in Step 1.1, design a truss structure model or a semi-rigid shell structure model to obtain the main load-bearing structure model of the UAV.

[0029] In specific exemplary technical solutions, truss structure models or semi-monocoque structure models can be designed based on design principles such as the shortest structural force transmission path, lightweight structure, and maintainability of airborne equipment. Among them, the design principle of the shortest force transmission path includes: the landing gear bow beam is directly connected to the fuselage frame beam to avoid direct diffusion and transmission of landing load / fuselage load; the design principle of lightweight structure includes: the tail section tube beam and fuselage skin and other load-bearing components are made of carbon fiber composite materials; the design principle of maintainability of airborne equipment includes: for example, the airborne equipment engine, the power unit and the airframe structure should have a certain gap (such as 25~50mm) to facilitate the maintenance of the power system.

[0030] Step 1.3: Based on the main load-bearing structure model of the UAV obtained in Step 1.2, the rods or flanges that transmit torque and axial force are abstracted into line models, the surfaces that transmit in-plane shear loads or out-of-plane loads are abstracted into surface models, the large mass blocks of the airborne system in the UAV are abstracted into mass points, and the corresponding connection relationships between them are established to obtain the wireframe skeleton model of the main load-bearing structure of the UAV.

[0031] Step 2: Based on the wireframe skeleton model obtained in Step 1, construct the initial finite element model of the main load-bearing structure of the UAV. During the construction of the initial finite element model, the main load-bearing components in the wireframe skeleton model that bear the longitudinal and lateral load transfer functions are geometrically modeled using beam elements (i.e., one-dimensional elements). The intermediate region formed by connecting and closing the aforementioned main load-bearing components is geometrically modeled using shell elements (two-dimensional elements). The initial finite element model is assigned the properties of a single selected material (generally, the material with the largest quantity used in the design, such as aluminum alloy 2A12, with a Young's modulus of 72 GPa and a Poisson's ratio of 0.33), and initial cross-sectional dimensions are set (interpretatively, the cross-sectional dimensions of a rod refer to its wall thickness and outer diameter, the cross-sectional dimensions of a flange refer to its wall thickness, and the cross-sectional dimensions of a panel refer to its wall thickness, etc.).

[0032] Step 3: Based on the selected multiple load conditions, load and perform finite element analysis on the initial finite element model constructed in Step 2 to obtain stress-displacement analysis results; based on the stress-displacement analysis results, iteratively adjust and optimize the initial cross-sectional dimensions so that the optimized cross-sectional dimensions meet all load conditions, and obtain the optimized finite element model.

[0033] In this step, the load conditions involved cover all expected operating states of the target structure on the ground and in the air, and the location, magnitude, and direction of the load are clearly defined for each condition, thus providing an input benchmark for finite element analysis. In the exemplary technical solution, various load conditions include: ground parking, ground gust, ground landing, ground mooring, ground hoisting, hovering, symmetrical maneuvering, discrete gust, fixed-point turning, and large-scale yaw load conditions.

[0034] In the technical solution of this invention embodiment, the principle for iteratively adjusting and optimizing the initial cross-sectional dimensions is as follows: When the force transmission structure is statically determinate, and the stress level in a local part of the component is significantly higher than that in other areas or exceeds the allowable value of the material design, its thickness or outer diameter should be increased until the stress level in the local part of the component is not much different from that in other areas. When the force transmission structure is statically indeterminate, and the stress level in a local part of the component is significantly higher than that in other areas or exceeds the allowable value of the material design, its thickness or outer diameter should be reduced until the stress level in the local part of the component is not much different from that in other areas. When the force transmission structure is statically determinate, and the stress level in a local area of ​​the component is significantly lower than that in other areas, its thickness or outer diameter is reduced until the stress level in the local area of ​​the component is not much different from that in other areas. When the force transmission structure is statically indeterminate, and the stress level in a local part of the component is significantly lower than that in other parts, its thickness or outer diameter is increased until the stress level in the local part of the component is not much different from that in other parts. At the same time, for locations with large component displacement, the functional requirements of the corresponding area should be met to prevent interference with the corresponding components. For example, when the components around the tail drive shaft undergo maximum deformation, they should not interfere with the tail drive shaft; when the mid-fuselage skin deforms, it should not interfere with the motion envelope of the servo motors running on the main reducer.

[0035] Step 4: Based on the optimized finite element model, the materials of each functional structural region are restored to the target materials to obtain a multi-material finite element model; based on the selected multiple working conditions, finite element analysis is performed on the multi-material finite element model to obtain stress-displacement analysis results, and adjustments are made based on the stress-displacement analysis results to obtain the adjusted multi-material finite element model (i.e., the design results of the main load-bearing structure of the UAV).

[0036] Specifically, for the optimized finite element model, based on the functional area division of the structure, the corresponding engineering material property parameters are retrieved from the material database and assigned to the corresponding element set of the model; at the same time, the equivalent replacement model of the cross-sectional dimensions is performed according to the principle of equivalent stiffness of the structure.

[0037] In a specific exemplary technical solution, the functional structural areas include: the forward fuselage, which needs to install equipment such as radar, is made of glass fiber composite material based on considerations of wave transmission and structural lightweighting; near the engine, titanium alloy TC4 material is selected for thermal insulation; the tail section tube beam transmits tail torque and torque, and carbon fiber composite tube beam is selected for structural lightweighting design; the longitudinal and transverse fuselage components are made of aluminum alloy 7050 / 7075 material based on considerations of structural lightweighting and manufacturability; the skid landing gear is made of stainless steel 30CrMnSiA based on considerations of low cost and manufacturing process; the fuselage skin is made of carbon fiber composite material based on considerations of structural lightweighting design; the tail boom tube adopts a honeycomb sandwich composite structure based on considerations of structural lightweighting design; and the horizontal tail adopts a foam sandwich composite structure based on considerations of structural lightweighting and manufacturability.

[0038] In a specific exemplary technical solution, adjustments are made based on the stress-displacement analysis results. Similar to the principle of iterative adjustment and optimization of the initial cross-sectional dimensions mentioned above, the allowable material design values ​​for different functional areas are optimized according to the actual assigned material properties at this stage. For example, the allowable stress for titanium alloy is 850 MPa, the allowable stress for alloy steel is 1000 MPa, the allowable tensile and compressive strain for carbon fiber composites is 3000~4000 με, and the allowable tensile and compressive strain for glass fiber composites is 3000~5000 με.

[0039] In a further preferred embodiment, after step 4, the following is also included: Step 5: Based on the adjusted multi-material finite element model, for specific key load-bearing components, such as pressure-bearing components like the main reducer mounting plate, heat-bearing components like the engine tail nozzle bracket, and fatigue-dependent components like the rotor shaft, the models are checked and adjusted according to criteria such as instability failure, thermal stress failure, and fatigue failure, to obtain the secondary adjusted multi-material finite element model as the design result of the main load-bearing structure of the UAV.

[0040] In the technical solution disclosed in the embodiments of the present invention, during the early design process of the main load-bearing structure of the UAV, based on its overall load conditions, airborne system layout and the functional and performance requirements of each station, three-dimensional modeling software (such as CATIA software) is used to perform parametric geometric modeling of the structural components using simplified one-dimensional beam elements and two-dimensional shell elements, thereby constructing its basic parametric mechanical model.

[0041] In the technical solution disclosed in the embodiments of the present invention, in the finite element analysis software, the line and surface models of the structural components are assigned a single material property and the initial cross-sectional dimensions are set. Specifically, high-strength aluminum alloy 2A12 can be initially selected as the structural material. Under this premise, the beam element cross-section is set to a circular tube with a diameter of 20mm and a wall thickness of 1.5mm, or a solid cross-section / standard profile with an equivalent thickness of 2mm is used; the shell element thickness is uniformly preset to 1mm.

[0042] In the technical solution disclosed in this invention embodiment, the allowable stress design criterion is applied according to different load conditions to iteratively adjust and optimize the cross-sectional dimensions until the structural strength design requirements are met. Specifically, based on the constructed parametric finite element model, all selected load conditions are applied to the main load-bearing structure of the UAV for mechanical analysis. According to the stress characteristics of components in different regions, the equivalent allowable stress criterion is used as the evaluation basis to implement bidirectional iterative optimization of stress and cross-sectional dimensions: when the stress of the component cross-section exceeds the allowable upper limit (>390MPa), its cross-sectional dimensions are appropriately increased; when the stress is lower than the allowable lower limit (<200MPa), the cross-sectional dimensions are reduced under the premise of meeting the minimum process constraints. Through the above iterative optimization strategy, the stress of each component under all load conditions meets the strength failure criterion and gradually converges to the optimal solution of cross-sectional dimensions that meets all load condition constraints.

[0043] In the technical solution disclosed in this invention, firstly, different material properties or composite material layup parameters are assigned to each functional structural region based on its functional requirements; then, according to the stiffness equivalence design principle, the cross-sectional dimensions are further adjusted until the allowable stress requirements of the materials in each region are met, while also taking into account other relevant performance indicators. Specifically, based on the finite element model that has achieved cross-sectional dimension convergence based on aluminum alloy material, and combined with the overall functional and performance requirements of the UAV structural system, the design and evaluation of multi-material schemes are further carried out: Firstly, the properties of metallic materials represented by titanium alloy and alloy steel, or the layup parameters and structural parameters corresponding to carbon fiber composites, glass fiber composites, foam sandwich composites, and honeycomb sandwich composites are assigned sequentially; then, based on the proportional relationship between the elastic modulus of the target material and the reference aluminum alloy, the converged cross-sectional dimensions are equivalently converted to obtain the initial dimensions of each material scheme; on this basis, the allowable stress intensity criterion of the corresponding material is used to iteratively optimize the cross-sectional dimensions.

[0044] In the technical solution disclosed in this invention embodiment, based on the above analysis, load data is extracted from some key components, and supplementary verification and analysis are performed using other applicable design criteria (such as fatigue strength, stability criteria, etc.) for specific working conditions. Specifically, based on the finite element model optimized by multi-material schemes, load data of key components are extracted, and extended strength and failure mode verification are further carried out according to relevant load conditions. The verification content includes criteria such as shear stability, compressive stability, and fatigue life to comprehensively evaluate the performance of the structure under complex loads. During the verification process, the analysis is strictly carried out according to the corresponding load specifications and usage requirements. If the component does not meet the requirements in terms of stability or fatigue life, its cross-sectional dimensions are appropriately increased; if the performance is significantly higher than the requirements and there is room for optimization, the cross-sectional dimensions are reduced under the premise of meeting process and functional constraints. Through the above iterative optimization process, the optimized design scheme that meets all strength, stability, and fatigue life requirements and has the smallest cross-sectional dimensions is finally obtained.

[0045] In a specific exemplary technical solution, in the structural design of a certain type of UAV, if the main load-bearing part adopts a truss structure, its materials may involve multiple choices such as aluminum alloy, titanium alloy, alloy steel, or carbon fiber composite materials. In the early stages of structural design, different cross-sectional types, dimensions, or ply parameters need to be set for the components; however, due to the large number of design variables, the iteration process often encounters problems such as concentrated computational resources and lengthy analysis cycles, leading to a significant decrease in the efficiency of structural finite element analysis, thus slowing down the overall design progress and restricting the deepening and optimization of subsequent solutions. The technical solution of this invention can improve the efficiency of defining the cross-sectional dimensions of UAV structures, thereby shortening the UAV development time.

[0046] In a specific exemplary technical solution of the present invention, a design method for the main load-bearing structure of an unmanned aerial vehicle (UAV) is provided, comprising the following steps: Step S1: In 3D modeling software, simplified line surfaces are used to perform geometric modeling of the UAV structural components to construct its basic mechanical model.

[0047] Step S2: In the finite element analysis software, assign the material properties of a single material, high-strength aluminum alloy 2A12, to the line and surface model of the structural component, and set a certain initial cross-sectional size.

[0048] Step S3: Based on different load conditions, apply the allowable stress design criteria to iteratively adjust and optimize the structural cross-sectional dimensions until the structural strength design requirements are met.

[0049] Step S4: Based on the above model and the functional requirements of each functional structural region, assign parameters to titanium alloy, alloy steel and T300 carbon fiber composite material respectively; according to the stiffness equivalence design principle, further adjust the cross-sectional dimensions until the allowable stress requirements of the materials in each region are met, while taking into account other relevant performance indicators. Step S5: Based on the above analysis, load data is extracted from some key components, and supplementary verification and analysis are performed using design criteria such as instability, thermal stress, or fatigue life for specific working conditions.

[0050] Traditional methods for designing lightweight multi-material main load-bearing structures for UAVs typically involve simultaneously optimizing cross-sectional dimensions using multiple materials. This leads to coupling between the cross-sectional parameters and strength criteria of different materials, resulting in numerous design variables, difficulty in iterative convergence, and low optimization efficiency when dealing with hundreds of load conditions. To address these issues, this invention proposes an integrated, hierarchical optimization strategy encompassing "material equivalence, dimension optimization, material restoration, and strength verification." It first uses a single type of material to optimize cross-sectional dimensions, then maps these dimensions to different actual materials for strength and failure mode verification. This decomposes the complex high-dimensional multi-material coupling problem into two sequential stages: single-material optimization and multi-material verification. This significantly reduces design complexity and iterative computation, and substantially improves optimization efficiency. The challenge in obtaining the improved methods of this invention lies in moving beyond the traditional mindset of "simultaneous material and dimension optimization." It requires recognizing that the dependence of dimension optimization on material properties can be decoupled. However, designers, limited by experience and conventional coupling analysis processes, often struggle to grasp this simplified modeling and step-by-step optimization approach of "unification first, mapping later." The technical solution disclosed in this invention simplifies the time required to define the cross-sectional dimensions of UAV structures and improves the iterative efficiency of structural design. Specifically, initially using a single material and applying allowable stress design criteria reduces the number of variables and avoids the complexity caused by differences in allowable stress and cross-sectional dimensions of various materials, thus enabling more efficient definition of cross-sectional dimensions. This invention can be applied to the field of aircraft and can also be extended to other corresponding mechanical fields.

[0051] The following are embodiments of the apparatus of the present invention, which can be used to execute embodiments of the method of the present invention. For details not disclosed in the apparatus embodiments, please refer to the embodiments of the method of the present invention.

[0052] Please see Figure 2 In this embodiment of the invention, a design system for the main load-bearing structure of an unmanned aerial vehicle (UAV) is provided, comprising: The wireframe skeleton model acquisition module is used to acquire the wireframe skeleton model of the main load-bearing structure of the UAV. The initial finite element model construction module is used to construct an initial finite element model of the main load-bearing structure of the UAV based on the wireframe skeleton model; wherein, in the initial finite element model, a single material property is assigned according to the material with the largest amount used in the UAV design, and the initial cross-sectional dimensions are set; The optimized finite element model acquisition module is used to load and perform finite element analysis on the initial finite element model based on selected multiple load conditions, obtain stress-displacement analysis results, and iteratively adjust and optimize the initial cross-sectional dimensions based on the stress-displacement analysis results so that the optimized cross-sectional dimensions meet all load conditions, thereby obtaining the optimized finite element model. The design result acquisition module is used to restore the materials of each functional structural region in the main load-bearing structure of the UAV to the target materials based on the optimized finite element model, and obtain a multi-material finite element model; based on the selected multiple load conditions, perform finite element analysis on the multi-material finite element model to obtain stress-displacement analysis results; adjust based on the stress-displacement analysis results; and use the adjusted multi-material finite element model as the design result of the main load-bearing structure of the UAV.

[0053] In one embodiment of the present invention, a computer device is provided, comprising a processor and a memory. The memory stores a computer program, which includes program instructions, and the processor executes the program instructions. The processor may be a Central Processing Unit (CPU), or other general-purpose processors, digital signal processors (DSPs), application-specific integrated circuits (ASICs), field-programmable gate arrays (FPGAs), or other programmable logic devices, discrete gate or transistor logic devices, discrete hardware components, etc. It is the computing and control core of the terminal, suitable for implementing one or more instructions, specifically suitable for loading and executing one or more instructions from a computer storage medium to achieve a corresponding method flow or corresponding function. The processor described in this embodiment of the present invention can be used to execute the design method of the main load-bearing structure of an unmanned aerial vehicle (UAV).

[0054] In one embodiment of the present invention, a storage medium is provided, specifically a computer-readable storage medium (Memory), which is a memory device in a computer device used to store programs and data. It is understood that the computer-readable storage medium here can include both the built-in storage medium in the computer device and extended storage media supported by the computer device. The computer-readable storage medium provides storage space that stores the operating system of the terminal. Furthermore, the storage space also stores one or more instructions suitable for loading and execution by a processor, which can be one or more computer programs (including program code). It should be noted that the computer-readable storage medium here can be high-speed RAM (Random Access Memory) or non-volatile memory, such as at least one disk storage device. The processor can load and execute one or more instructions stored in the computer-readable storage medium to implement the corresponding steps of the design method for the main load-bearing structure of the UAV in the above embodiments.

[0055] Those skilled in the art will understand that embodiments of the present invention can be provided as methods, systems, or computer program products. Therefore, the present invention can take the form of a completely hardware embodiment, a completely software embodiment, or an embodiment combining software and hardware aspects. Furthermore, the present invention can take the form of a computer program product embodied on one or more computer-usable storage media (including, but not limited to, disk storage, optical storage, etc.) containing computer-usable program code.

[0056] This invention is described with reference to flowchart illustrations and / or block diagrams of methods, apparatus (systems), and computer program products according to embodiments of the invention. It will be understood that each block of the flowchart illustrations and / or block diagrams, and combinations of blocks in the flowchart illustrations and / or block diagrams, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, special-purpose computer, embedded processor, or other programmable data processing apparatus to produce a machine, such that the instructions, which execute via the processor of the computer or other programmable data processing apparatus, generate instructions for implementing the flowchart illustrations and / or block diagrams. Figure 1 One or more processes and / or boxes Figure 1 A device that provides the functions specified in one or more boxes.

[0057] These computer program instructions may also be stored in a computer-readable storage medium that can direct a computer or other programmable data processing device to function in a particular manner, such that the instructions stored in the computer-readable storage medium produce an article of manufacture including instruction means, which are implemented in a process Figure 1One or more processes and / or boxes Figure 1 The function specified in one or more boxes.

[0058] These computer program instructions may also be loaded onto a computer or other programmable data processing equipment to cause a series of operational steps to be performed on the computer or other programmable equipment to produce a computer-implemented process, thereby providing instructions that execute on the computer or other programmable equipment for implementing the process. Figure 1 One or more processes and / or boxes Figure 1 The steps of the function specified in one or more boxes.

[0059] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention and not to limit it. Although the present invention has been described in detail with reference to the above embodiments, those skilled in the art should understand that modifications or equivalent substitutions can still be made to the specific implementation of the present invention. Any modifications or equivalent substitutions that do not depart from the spirit and scope of the present invention should be covered within the scope of protection of the claims of the present invention.

Claims

1. A design method of a main load-bearing structure of a drone, characterized in that, Includes the following steps: Obtain the wireframe skeleton model of the main load-bearing structure of the UAV; Based on the wireframe skeleton model, an initial finite element model of the main load-bearing structure of the UAV is constructed; wherein, in the initial finite element model, a single material property is assigned according to the material with the largest amount used in the UAV design, and the initial cross-sectional dimensions are set; Based on the selected multiple load conditions, the initial finite element model is loaded and subjected to finite element analysis to obtain stress-displacement analysis results. Based on the stress-displacement analysis results, the initial cross-sectional dimensions are iteratively adjusted and optimized so that the optimized cross-sectional dimensions meet all load conditions, and the optimized finite element model is obtained. Based on the optimized finite element model, the materials of each functional structural region in the main load-bearing structure of the UAV are restored to the target materials to obtain a multi-material finite element model. Based on the selected multiple load conditions, finite element analysis is performed on the multi-material finite element model to obtain stress-displacement analysis results. Based on the stress-displacement analysis results, adjustments are made, and the adjusted multi-material finite element model is used as the design result of the main load-bearing structure of the UAV.

2. The method of claim 1, wherein, The specific steps for obtaining the wireframe skeleton model of the main load-bearing structure of the UAV include: Obtain the design requirements for the main load-bearing structure of the UAV; wherein, the design requirements include: overall load conditions and airborne system layout; Based on the design requirements, design a truss structure model or a semi-rigid shell structure model to obtain the main load-bearing structure model of the UAV. Based on the main load-bearing structure model of the UAV, wireframe modeling is performed to obtain the wireframe skeleton model of the main load-bearing structure of the UAV. In the process of wireframe modeling, according to the overall load conditions and the layout of the airborne system, the rods or flanges that transmit torque and axial force are abstracted into line models, the surfaces that transmit in-plane shear loads or out-of-plane loads are abstracted into surface models, and the large mass blocks of the airborne system are abstracted into mass points. The connection relationship between the line models, surface models and mass points is established.

3. The method of claim 2, wherein, In the process of designing a truss structure model or a semi-rigid shell structure model based on the aforementioned design requirements to obtain the main load-bearing structure model of the UAV, the design principles of shortest structural force transmission path, lightweight structure, and maintainability of airborne equipment are adopted when designing the truss structure model or the semi-rigid shell structure model.

4. The method of claim 1, wherein, In the process of constructing the initial finite element model of the main load-bearing structure of the UAV based on the wireframe skeleton model, the main load-bearing components that bear the longitudinal and lateral load transfer functions in the wireframe skeleton model are geometrically modeled using beam elements; the intermediate region formed by connecting and closing the main load-bearing components is geometrically modeled using shell elements.

5. The design method for the main load-bearing structure of an unmanned aerial vehicle according to claim 1, characterized in that, The initial cross-sectional dimensions are iteratively adjusted and optimized based on the stress-displacement analysis results to ensure that the optimized cross-sectional dimensions meet all load conditions. The principles for iteratively adjusting and optimizing the initial cross-sectional dimensions during the process of obtaining the optimized finite element model include: The force transmission structure is statically determinate. When the stress level in a local area of ​​the component is higher than that in other areas and the difference is greater than or equal to the first preset difference threshold, or when the stress level in a local area of ​​the component is greater than or equal to the allowable value of the material design, the thickness or outer diameter of the component is increased until the local stress level of the component meets the requirements. When the stress level in a local area of ​​the component is lower than that in other areas and the difference is greater than or equal to the second preset difference threshold, the thickness or outer diameter of the component is reduced until the local stress level of the component meets the requirements. The force transmission structure is statically indeterminate. When the stress level in a local area of ​​the component is higher than that in other areas and the difference is greater than or equal to the third preset difference threshold, the component thickness or outer diameter is reduced until the local stress level of the component meets the requirements. When the stress level in a local area of ​​the component is lower than that in other areas and the difference is greater than or equal to the fourth preset difference threshold, the component thickness or outer diameter is increased until the local stress level of the component meets the requirements. Among them, the components meet the functional requirements of the corresponding functional area at the position of maximum displacement, and there is no interference between the functional components.

6. The design method for the main load-bearing structure of an unmanned aerial vehicle according to claim 1, characterized in that, Based on the optimized finite element model, the steps to restore the material of each functional structural region in the main load-bearing structure of the UAV to the target material and obtain the multi-material finite element model include: Based on the division of each functional structural region, the property parameters of the target material are called and assigned to the corresponding unit set; Based on the principle of equivalent stiffness of structure, the cross-sectional dimensions of the element set are replaced by equivalent modeling to obtain a multi-material finite element model.

7. The design method for the main load-bearing structure of an unmanned aerial vehicle according to claim 1, characterized in that, Based on selected multiple load conditions, finite element analysis is performed on the multi-material finite element model to obtain stress-displacement analysis results. Adjustments are made based on these results, and the adjusted multi-material finite element model is used as the design result for the main load-bearing structure of the UAV. After adjustments based on the stress-displacement analysis results, the following verification and adjustment steps are also included: The pressure-bearing components were checked and adjusted according to the instability failure criteria; The heated components were checked and adjusted according to the thermal stress failure criteria; Fatigue-dependent components should be checked and adjusted according to fatigue failure criteria.

8. A design system for the main load-bearing structure of an unmanned aerial vehicle (UAV), characterized in that, include: The wireframe skeleton model acquisition module is used to acquire the wireframe skeleton model of the main load-bearing structure of the UAV. The initial finite element model construction module is used to construct an initial finite element model of the main load-bearing structure of the UAV based on the wireframe skeleton model; wherein, in the initial finite element model, a single material property is assigned according to the material with the largest amount used in the UAV design, and the initial cross-sectional dimensions are set; The optimized finite element model acquisition module is used to load and perform finite element analysis on the initial finite element model based on selected multiple load conditions, obtain stress-displacement analysis results, and iteratively adjust and optimize the initial cross-sectional dimensions based on the stress-displacement analysis results so that the optimized cross-sectional dimensions meet all load conditions, thereby obtaining the optimized finite element model. The design result acquisition module is used to restore the materials of each functional structural region in the main load-bearing structure of the UAV to the target materials based on the optimized finite element model, and obtain a multi-material finite element model; based on the selected multiple load conditions, perform finite element analysis on the multi-material finite element model to obtain stress-displacement analysis results; adjust based on the stress-displacement analysis results; and use the adjusted multi-material finite element model as the design result of the main load-bearing structure of the UAV.

9. An electronic device comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that, When the processor executes the computer program, it implements the design method for the main load-bearing structure of the UAV as described in any one of claims 1 to 7.

10. A non-transitory computer-readable storage medium having a computer program stored thereon, characterized in that, When the computer program is executed by the processor, it implements the design method for the main load-bearing structure of the UAV as described in any one of claims 1 to 7.

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