A large-load lightweight unmanned aerial vehicle platform design method

Abstract

The present application relates to the technical field of unmanned aerial vehicle design, in particular to a large-load lightweight unmanned aerial vehicle platform design method, comprising: obtaining task profile parameters such as maximum take-off weight and payload mass, constructing an initial overall layout scheme based on the parameters and performing mass characteristic calculation to obtain an initial mass characteristic data set; inputting the data set into a structure topology optimization model, taking minimization of the overall machine empty weight as the target, taking the payload bending moment and the maximum take-off weight axial force as the constraints, and generating a main load-bearing structure transmission path distribution diagram; optimizing material distribution according to the diagram to obtain a mixed material scheme of high-stiffness composite material and local reinforced metal parts; converting the scheme into detailed structure design drawings and preparing manufacturing process cards. The method can accurately optimize the transmission path and material distribution, balance the lightweight and large-load carrying capacity, reduce the waste of invalid structure and material, and improve the structural stability and manufacturing precision.

Description

Technical Field

[0001] This invention relates to the field of unmanned aerial vehicle (UAV) design technology, and in particular to a design method for a lightweight UAV platform with a large payload. Background Technology

[0002] Heavy-payload lightweight drone platforms are widely used in logistics transportation, emergency rescue, aerial surveying and mapping and other fields. The core design of these drones is to balance the heavy-payload capacity with the lightweight requirement. In the existing technology, the initial overall layout of the drone is usually determined according to the mission requirements. Then, the shape and size of the main load-bearing structure are determined by experience or simple structural calculations. Subsequently, a single type of material is selected for structural manufacturing. Some designs may use topology optimization technology, but most of them aim to minimize stress or use only a single load as a constraint.

[0003] Existing technical solutions have shortcomings. The initial overall layout design did not fully incorporate mission profile parameters for mass characteristic calculations, resulting in a lack of accurate data support for subsequent structural optimization. The topology optimization process did not comprehensively consider the bending moment generated by the payload mass and the axial force generated by the maximum takeoff weight, leading to an unreasonable force transmission path in the generated main load-bearing structure and making it difficult to minimize the overall empty weight. In terms of material selection, while single composite materials can achieve lightweighting, their local load-bearing capacity is insufficient. While single metal materials have high strength, they significantly increase the overall weight. The existing hybrid material distribution design is not optimized according to the force transmission path, making it impossible to achieve precise material matching. This makes it difficult for the UAV to meet the requirements of large payload, lightweighting, and structural strength. At the same time, the manufacturing process is not precisely adapted to the material distribution scheme, affecting the performance of the structure. Summary of the Invention

[0004] The purpose of this invention is to address the shortcomings of existing technologies by proposing a design method for a lightweight unmanned aerial vehicle (UAV) platform with a large payload.

[0005] To achieve the above objectives, the present invention adopts the following technical solution: a design method for a heavy-payload lightweight unmanned aerial vehicle platform, comprising:

[0006] Obtain the mission profile parameters of the large-payload lightweight unmanned aerial vehicle platform, including the maximum takeoff weight, payload mass, endurance index, and typical flight speed range;

[0007] Based on the mission profile parameters, an initial overall layout scheme for the heavy-payload lightweight UAV platform is constructed.

[0008] The initial overall layout scheme is subjected to mass characteristic calculation to obtain an initial mass characteristic dataset containing the center of gravity coordinates, rotational inertia matrix and the empty weight of the whole machine;

[0009] The initial mass characteristic dataset is input into the structural topology optimization model. With the goal of minimizing the empty weight of the entire aircraft and the constraints of the bending moment generated by the payload mass and the axial force generated by the maximum takeoff weight, a force transmission path distribution diagram of the main load-bearing structure is generated.

[0010] Based on the force transmission path distribution diagram, the material distribution of the main load-bearing structure is optimized to obtain a hybrid material distribution scheme consisting of a high-stiffness composite material region and locally reinforced metal parts;

[0011] The hybrid material distribution scheme is converted into detailed structural design drawings of the heavy-load lightweight UAV platform. Based on the detailed structural design drawings, a manufacturing process card for the heavy-load lightweight UAV platform is prepared. The manufacturing process card specifies the curing temperature of the composite material and the machining tolerances of the metal parts.

[0012] As a further aspect of the present invention, the initial overall layout scheme is subjected to mass characteristic calculation to obtain an initial mass characteristic dataset containing the center of gravity coordinates, rotational inertia matrix, and overall empty weight, including:

[0013] The initial overall layout scheme defines the fuselage aspect ratio, wing aspect ratio, and power system installation location;

[0014] In the initial overall layout scheme, geometric modeling is performed on each component of the airframe to generate a geometric model that includes the fuselage outer curved surface, the wing planar shape, and the landing gear strut positions;

[0015] Assign corresponding material density properties to each part of the geometric model, wherein the material density properties are derived from a preset material database;

[0016] By discretizing the geometric model, the centroid position and volume of each discrete unit are calculated, and the centroid position and volume data of all discrete units are summarized.

[0017] Based on the centroid position and volume data of all discrete units, the centroid coordinates and empty weight of the entire machine of the initial overall layout scheme are calculated.

[0018] Based on the mass distribution of the geometric model and the spatial position of the discrete units, the rotational inertia components around the three main axes of the body are calculated and assembled to form the rotational inertia matrix.

[0019] The center of gravity coordinates, moment of inertia matrix, and empty weight of the entire machine are packaged and encapsulated to form the initial mass characteristic dataset.

[0020] As a further aspect of the present invention, the initial mass characteristic dataset is input into a structural topology optimization model. With the goal of minimizing the empty weight of the entire aircraft and constrained by the bending moment generated by the payload mass and the axial force generated by the maximum takeoff weight, a force transmission path distribution diagram of the main load-bearing structure is generated, including:

[0021] The empty weight of the aircraft is extracted from the initial mass characteristic dataset as the benchmark value for optimization, and the bending moment generated by the payload mass and the axial force generated by the maximum takeoff weight are extracted as the working condition load boundary.

[0022] The working condition load boundary and the benchmark value are imported into the structural topology optimization model, and the number of optimization iterations is set to a preset fixed value;

[0023] In the structural topology optimization model, the empty weight of the whole machine is taken as the objective function, and the structural stiffness under the working condition load boundary is taken as the constraint condition to start the iterative calculation process;

[0024] In each iterative calculation, the material's state of existence in space is adjusted based on the current stress distribution cloud map and strain energy density distribution cloud map;

[0025] When the number of iterations reaches a preset fixed value, the iteration process is terminated, and the boundary contours of the high-density and low-density regions of the material at this time are extracted.

[0026] Map the boundary contour back to the body coordinate system of the heavy-load lightweight UAV platform and draw the force transmission path distribution diagram of the main load-bearing structure;

[0027] The steps for constructing the structural topology optimization model include:

[0028] Obtain a three-dimensional geometric model of the initial overall layout scheme of the large-payload lightweight UAV platform, wherein the three-dimensional geometric model includes the original design space of the main load-bearing structure;

[0029] Material properties are applied to the three-dimensional geometric model, and the entire design space is discretized into a uniformly distributed finite element mesh;

[0030] Multiple load application points and displacement constraint points are defined on the boundary of the design space. The load application points correspond to the working load boundaries in the initial mass characteristic dataset, and the displacement constraint points correspond to the fixing methods of the landing gear connection points and the power system mounting points.

[0031] The objective function of the structural topology optimization model is set as minimizing the overall empty weight of the machine;

[0032] Multiple optimization constraints are input into the structural topology optimization model. The optimization constraints include that the maximum structural deformation under the bending moment generated by the effective payload mass does not exceed a preset value, and the maximum structural stress under the axial force generated by the maximum takeoff weight does not exceed the allowable stress of the material.

[0033] Define the iterative convergence criterion for the structural topology optimization model, wherein the iterative convergence criterion is that the change in the material volume fraction of the design space calculated in two adjacent iterations is less than a preset threshold.

[0034] As a further aspect of the present invention, based on the force transmission path distribution diagram, the material distribution of the main load-bearing structure is optimized to obtain a hybrid material distribution scheme consisting of a high-stiffness composite material region and locally reinforced metal parts, including:

[0035] The high-density areas in the force transmission path distribution map are marked as critical load-bearing areas, and the low-density areas are marked as secondary load-bearing areas.

[0036] For the critical load-bearing area, carbon fiber composite material grades with tensile strength and compressive strength both higher than the threshold are selected from the preset material database and identified as the high stiffness composite material region;

[0037] For the secondary load-bearing area, the original lightweight foam filling structure is retained to maintain the buoyancy characteristics and low-speed aerodynamic efficiency of the body;

[0038] In the critical load-bearing area, the joint locations bearing concentrated loads are identified, and titanium alloys or aluminum alloys with high yield strength are selected from the material database as the local reinforcing metal parts.

[0039] The determined positions of the high-rigidity composite material region, the lightweight foam filling structure, and the locally reinforcing metal parts within the body structure are combined to form the hybrid material distribution scheme.

[0040] As a further aspect of the present invention, the hybrid material distribution scheme is transformed into detailed structural design drawings of the heavy-load lightweight unmanned aerial vehicle platform, including:

[0041] The detailed structural design drawings include skin ply angles and detailed drawings of the connection nodes of the internal skeleton;

[0042] Read the material type identifiers and spatial location information in the hybrid material distribution scheme, and establish a one-to-one correspondence table between materials and structural components;

[0043] Based on the one-to-one correspondence table, the skin of the heavy-load lightweight UAV platform is parametrically modeled in the three-dimensional design environment. The inputs for the parametric modeling include the layup thickness and fiber orientation angle of the high-stiffness composite material.

[0044] The internal skeleton of the heavy-load lightweight UAV platform is modeled as a rod structure, and in the rod structure modeling, the local reinforcing metal parts are defined as rigid connection nodes;

[0045] Based on the parametric modeling and the rod structure modeling, an assembly interference check is performed to correct structural conflicts caused by changes in material thickness.

[0046] The 3D model, after interference checking and correction, is projected onto a 2D engineering drawing, and the skin ply angle and the connection node details of the internal skeleton are marked to form the detailed structural design drawings.

[0047] As a further aspect of the present invention, based on the detailed structural design drawings, a manufacturing process card for the heavy-payload lightweight unmanned aerial vehicle platform is prepared, including:

[0048] The names of all components made of the high-rigidity composite material and their ply structures are extracted from the detailed structural design drawings.

[0049] For each component made of the high-rigidity composite material, an autoclave molding process route is formulated, in which the heating rate, holding time and cooling rate are set as control parameters for the curing temperature of the composite material.

[0050] The dimensions of all parts composed of the locally reinforcing metal parts are extracted from the detailed structural design drawings, and their nominal dimensions and tolerance zones are calculated.

[0051] For each part composed of the aforementioned locally reinforcing metal parts, a CNC machining center is selected as the machining equipment. The calculated tolerance zone is converted into the machining tolerance of the metal parts, and a machining program is compiled.

[0052] The autoclave forming process route and the processing procedure are integrated into a manufacturing process card for the heavy-load lightweight UAV platform.

[0053] As a further aspect of the present invention, after constructing the initial overall layout scheme of the heavy-payload lightweight UAV platform based on the mission profile parameters, the invention further includes:

[0054] The initial overall layout scheme is imported into the aerodynamic interference analysis module to simulate the disturbance of the wing flow field after the payload is mounted.

[0055] The pressure coefficient distribution data output by the aerodynamic interference analysis module is collected to identify high-risk areas of airflow separation on the wing surface;

[0056] Based on the location of the high-risk area of ​​airflow separation, the installation position and deflection angle of the wing trailing edge flaps are adjusted in the initial overall layout scheme;

[0057] The adjusted parameters of the wing trailing edge flaps are fed back to the initial overall layout scheme, updating the geometric configuration data of the initial overall layout scheme.

[0058] As a further aspect of the present invention, the adjusted parameters of the wing trailing edge flaps are fed back to the initial overall layout scheme to update the geometric configuration data of the initial overall layout scheme, including:

[0059] Read the adjustment parameters of the trailing edge flap of the wing, including the deflection angle value and the offset distance relative to the wing chord line;

[0060] In the geometric model of the initial overall layout scheme, locate the coordinates of the drive hinge point of the trailing edge flap of the wing;

[0061] Based on the deflection angle and the offset distance, the new coordinates of the end profile point of the wing trailing edge flap in the body coordinate system are recalculated.

[0062] The three-dimensional solid model of the wing trailing edge flap is reconstructed by replacing the old trailing edge point coordinates in the geometric model with the new trailing edge point coordinates.

[0063] The reconstructed 3D solid model of the wing trailing edge flap replaces the original wing trailing edge flap model in the initial overall layout scheme, thus completing the update of the geometric configuration data.

[0064] As a further aspect of the present invention, before converting the hybrid material distribution scheme into detailed structural design drawings of the heavy-load lightweight UAV platform, the invention further includes:

[0065] Buckling stability is checked in the high-stiffness composite material region of the hybrid material distribution scheme, and the load data used for buckling stability check is the empty weight of the whole machine in the initial mass characteristic dataset;

[0066] If the verification result does not meet the preset stability margin requirement, the number of local layups in the high stiffness composite material region is increased, and the buckling stability verification is performed again.

[0067] When the verification result meets the stability margin requirement, the number of layups in the high-stiffness composite material region at this time is confirmed as the final design value.

[0068] Substitute the high-stiffness composite material region data containing the final design value into the hybrid material distribution scheme, and then perform the step of converting it into detailed structural design drawings.

[0069] As a further aspect of the present invention, after compiling the manufacturing process card for the large-payload lightweight unmanned aerial vehicle platform, the method further includes:

[0070] Extract the curing temperature curve of the composite material and the machining tolerance range of the metal part from the manufacturing process card to form quantifiable quality control indicators;

[0071] The quality control indicators are entered into the production quality control system of the heavy-payload lightweight UAV platform as the inspection benchmark for subsequent trial production batches.

[0072] During the production process, the curing temperature monitoring data and machining dimension detection data of the actual prototype parts are collected and compared with the recorded quality control indicators;

[0073] The deviation values ​​in the comparison results are fed back to the structural topology optimization model as input data for adjusting constraints in the next design iteration.

[0074] Compared with the prior art, the advantages and positive effects of the present invention are as follows:

[0075] The initial mass characteristic dataset is input into the structural topology optimization model. With the goal of minimizing the overall empty weight and constrained by the bending moment generated by the payload mass and the axial force generated by the maximum takeoff weight, a force transmission path distribution diagram of the main load-bearing structure is generated. This scheme makes the topology optimization process more closely resemble the actual working conditions of the UAV. The design of the force transmission path is based on accurate mass characteristic data and comprehensive load constraints. Compared to conventional topology optimization methods that only focus on a single objective or constraint, it can accurately locate the key stress areas of the main load-bearing structure, reduce ineffective structural design, and reduce the overall empty weight while meeting the requirements for high load bearing capacity. This allows the UAV to carry a larger payload or extend its endurance under the same power conditions, while also optimizing the structural stress state and reducing stress concentration.

[0076] Based on the force transmission path distribution diagram, the material distribution of the main load-bearing structure is optimized, resulting in a hybrid material distribution scheme consisting of high-stiffness composite material regions and locally reinforced metal components. This scheme achieves precise matching between the materials and the force transmission path of the main load-bearing structure. The high-stiffness composite material regions are used to bear dispersed conventional loads, fully leveraging the advantages of lightweight and high stiffness of composite materials to reduce the overall weight. The locally reinforced metal components are used to bear concentrated loads on the force transmission path, compensating for the insufficient local load-bearing capacity of composite materials. Compared with conventional single-material structures or disordered hybrid material structures, this scheme avoids both the problem of insufficient local strength of single composite materials and the drawback of excessive weight of single metal materials, achieving a balance between lightweight and structural strength. At the same time, the reasonable material distribution can reduce material waste and lower manufacturing costs. With the corresponding manufacturing process cards, it can ensure that the material performance is fully utilized, improving the manufacturing accuracy and stability of the structure. Attached Figure Description

[0077] Figure 1 This is a flowchart of a design method for a large-payload lightweight unmanned aerial vehicle platform according to the present invention;

[0078] Figure 2 The curves showing the changes in relative empty weight ratio and relative stiffness ratio during the iterative process of structural topology optimization;

[0079] Figure 3 A flowchart for converting a mixed material distribution scheme into detailed structural design drawings;

[0080] Figure 4 This is a curve showing the relationship between flap deflection angle and surface pressure coefficient.

[0081] Figure 5 The curves showing the variation of curing temperature deviation and processing tolerance deviation with the trial batch are shown. Detailed Implementation

[0082] To make the objectives, technical solutions, and advantages of this invention clearer, the invention will be further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative and not intended to limit the invention.

[0083] In the description of this invention, it should be understood that the terms "length," "width," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," and "outer," etc., indicating orientation or positional relationships, are based on the orientation or positional relationships shown in the accompanying drawings and are only for the convenience of describing the invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation, and therefore should not be construed as a limitation of the invention. Furthermore, in the description of this invention, "a plurality of" means two or more, unless otherwise explicitly specified.

[0084] See Figure 1 The mission profile parameters of the UAV platform were obtained, including maximum takeoff weight, payload mass, endurance, and typical flight speed range. Based on these parameters, an initial overall layout scheme for the UAV platform was constructed. Mass characteristics of this initial layout scheme were calculated, resulting in an initial mass characteristic dataset containing the center of gravity coordinates, moment of inertia matrix, and empty weight. This dataset was then input into a pre-defined structural topology optimization model. The model's optimization objective was to minimize the empty weight, using the bending moment generated by the payload mass and the axial force generated by the maximum takeoff weight as constraints. The model calculated and generated a force transmission path distribution diagram for the main load-bearing structure. Based on this diagram, material distribution optimization was performed on the main load-bearing structure, resulting in a hybrid material distribution scheme consisting of high-stiffness composite material regions and locally reinforced metal components. Finally, the hybrid material distribution scheme was transformed into detailed structural design drawings for the UAV platform. Based on these drawings, a manufacturing process card specifying the curing temperature of the composite material and the machining tolerances of the metal components was developed.

[0085] In one embodiment of the present invention, after obtaining the mission profile parameters and constructing an initial overall layout scheme, its mass characteristics are calculated to obtain an initial mass characteristic dataset. The initial overall layout scheme defines the fuselage slenderness ratio, wing aspect ratio, and power system installation location. In this scheme, geometric modeling is performed on each component of the fuselage to generate a geometric model including the fuselage outer surface, wing planar shape, and landing gear strut positions. Corresponding material density properties are assigned to each part of this geometric model, and these material density properties are derived from a preset material database. By discretizing the geometric model, the center of mass position and volume of each discrete unit are calculated, and the center of mass position and volume data of all discrete units are summarized. Based on the center of mass position and volume data of all discrete units, the center of gravity coordinates and the empty weight of the aircraft in the initial overall layout scheme are calculated. According to the mass distribution of the geometric model and the spatial position of the discrete units, the rotational inertia components around the three principal axes of the fuselage are calculated and assembled to form a rotational inertia matrix. The calculated center of gravity coordinates, rotational inertia matrix, and empty weight of the aircraft are packaged and encapsulated to form the initial mass characteristic dataset.

[0086] In practical implementation, the initial overall layout scheme defines the fuselage slenderness ratio, wing aspect ratio, and power system installation location. Within this initial scheme, geometric modeling is performed on each component of the fuselage, generating a geometric model that includes the fuselage's curved surfaces, wing planar shape, and landing gear strut positions. Each part of the geometric model is assigned a corresponding material density attribute, derived from a pre-set material database. The geometric model is discretized, and the center of mass position and volume of each discrete unit are calculated, summing the data for all discrete units. In practical implementation, based on the center of mass position and volume data of all discrete units, the coordinates of the center of gravity and the empty weight of the aircraft in the initial overall layout scheme are calculated. This calculation involves summing the contribution of each discrete unit. It can be understood that, based on the mass distribution of the geometric model and the spatial positions of the discrete units, the rotational inertia components around the three principal axes of the fuselage are calculated and assembled to form the rotational inertia matrix. One way to express the calculated center of gravity coordinates is as follows:

[0087] ;

[0088] ;

[0089] ;

[0090] in: These represent the coordinates of the center of gravity along the three axes of the body coordinate system. This represents the total number of discrete units. Indicates the first The material density assigned to each discrete unit Indicates the first The volume of a discrete unit Indicates the first The coordinates of the centroid of each discrete unit in the body coordinate system. This can be understood as packaging the calculated centroid coordinates, moment of inertia matrix, and total empty weight of the machine into an initial mass characteristic dataset.

[0091] In one embodiment of the present invention, a three-dimensional geometric model of the initial overall layout scheme of a heavy-load lightweight unmanned aerial vehicle (UAV) platform is obtained. This three-dimensional geometric model includes the original design space of the main load-bearing structure. Material properties are applied to this three-dimensional geometric model, and the entire design space is discretized into a uniformly distributed finite element mesh. Multiple load application points and displacement constraint points are defined on the boundary of the design space. The load application points correspond to the working load boundaries in the initial mass characteristic dataset, and the displacement constraint points correspond to the fixing methods of the landing gear connection points and the power system mounting points. The objective function of the structural topology optimization model is set as minimizing the empty weight of the entire aircraft. Multiple optimization constraints are input into the model, including that the maximum structural deformation under the bending moment generated by the payload mass does not exceed a preset value, and the maximum structural stress under the axial force generated by the maximum takeoff weight does not exceed the allowable stress of the material. An iterative convergence criterion for the model is defined, which is that the change in the material volume fraction of the design space calculated in two adjacent iterations is less than a preset threshold. In application, the empty weight of the entire aircraft is extracted from the initial mass characteristic dataset as the benchmark value for optimization, and the bending moment generated by the payload mass and the axial force generated by the maximum takeoff weight are extracted as the working load boundaries. The load boundary and baseline values ​​are imported into the structural topology optimization model, and the number of optimization iterations is set to a preset fixed value. In the model, the empty weight of the entire aircraft is used as the objective function, and the structural stiffness under the load boundary is used as the constraint condition to initiate the iterative calculation process. In each iteration, the spatial state of the material is adjusted based on the current stress distribution cloud map and strain energy density distribution cloud map. The iteration terminates when the number of iterations reaches the preset fixed value, and the boundary contours of the high-density and low-density regions of the material are extracted. This boundary contour is mapped back to the body coordinate system of the UAV platform to draw the force transmission path distribution diagram of the main load-bearing structure.

[0092] See Figure 2In the topology optimization iteration process of the main load-bearing structure of the lightweight UAV platform, the changing trends of the relative empty weight ratio and the relative stiffness ratio intuitively reflect the synergistic optimization effect of lightweighting and stiffness improvement. In the figure, the blue curve represents the relative empty weight ratio based on the initial scheme, and the orange curve represents the relative stiffness ratio based on the initial scheme. As the number of iterations increases from 0 to 20, the relative empty weight ratio continuously decreases monotonically from the initial 0.96 to 0.20, achieving a significant lightweighting target, while the relative stiffness ratio continuously increases monotonically from the initial 0.81 to 1.00, completing the strengthening of structural stiffness while achieving lightweighting. The two curves show a typical optimization convergence characteristic of one increasing and the other decreasing, verifying the effectiveness of the topology optimization model with the goal of minimizing the overall empty weight and the constraint of structural stiffness, and providing a quantitative basis for determining the optimal force transmission path and material distribution scheme.

[0093] Based on the force transmission path distribution diagram, the material distribution of the main load-bearing structure is optimized. High-density areas in the force transmission path distribution diagram are marked as critical load-bearing areas, and low-density areas are marked as secondary load-bearing areas. For the critical load-bearing areas, carbon fiber composite materials with tensile and compressive strengths exceeding thresholds are selected from a pre-set material database and identified as high-stiffness composite material regions. For the secondary load-bearing areas, the original lightweight foam-filled structure is retained to maintain the buoyancy characteristics and low-speed aerodynamic efficiency of the aircraft. Within the critical load-bearing areas, the joint locations bearing concentrated loads are identified, and high-yield-strength titanium alloys or aluminum alloys are selected from the material database as local reinforcing metal components. The determined high-stiffness composite material regions, lightweight foam-filled structures, and local reinforcing metal components are combined within the aircraft structure to form a hybrid material distribution scheme.

[0094] In practical implementation, the construction of the structural topology optimization model involves obtaining a three-dimensional geometric model of the initial overall layout scheme of the heavy-load lightweight UAV platform. This three-dimensional geometric model contains the original design space of the main load-bearing structure. Isotropic material properties are applied to the three-dimensional geometric model, and the entire design space is discretized into a uniformly distributed hexahedral finite element mesh. Load conditions corresponding to the bending moment generated by the effective payload mass and the axial force generated by the maximum takeoff weight are defined at the load application points in the design space. Displacement constraint points correspond to the fixing methods of the landing gear connection points and the power system mounting points. The objective function of the structural topology optimization model is set as minimizing the empty weight of the entire aircraft. Multiple optimization constraints are input into the structural topology optimization model, including ensuring that the maximum structural deformation under the bending moment generated by the effective payload mass does not exceed a preset value, and that the maximum structural stress under the axial force generated by the maximum takeoff weight does not exceed the allowable stress of the material. It can be understood that an iterative convergence criterion for the structural topology optimization model is defined: the change in the material volume fraction of the design space obtained from two adjacent iterations is less than 0.05%.

[0095] In the application of structural topology optimization models, the empty weight of the entire aircraft is extracted from the initial mass characteristic dataset as the benchmark value for optimization, and the bending moment generated by the payload mass and the axial force generated by the maximum takeoff weight are extracted as the working condition load boundaries. In some embodiments, the working condition load boundaries and benchmark values ​​are imported into the structural topology optimization model, and the number of optimization iterations is set to two hundred. In the structural topology optimization model, the empty weight of the entire aircraft is used as the objective function, and the structural stiffness under the working condition load boundaries is used as the constraint condition to start the iterative calculation process. In each iteration, the existence state of the material in space is adjusted according to the current stress distribution cloud map and strain energy density distribution cloud map using the optimization criterion method. When the number of iterations reaches two hundred, the iteration process is terminated, and the boundary contours of the high-density and low-density regions of the material at this time are extracted. Optionally, the boundary contours are mapped back to the body coordinate system of the heavy-load lightweight UAV platform to draw the force transmission path distribution diagram of the main load-bearing structure. Among them, a mathematical expression of the objective function of minimizing the empty weight of the entire aircraft involves summing the material contributions of all elements in the design domain:

[0096] ;

[0097] Where: symbol This represents the estimated empty weight of the entire machine through optimization iteration, with the symbol... Represents the total number of finite elements after discretization in the design space, symbol It is the unit's index number, symbol Indicates the first The relative density design variable for each unit varies continuously between 0 and 1, with the sign... Indicates a The interpolation function is used to correlate the relative density of an element with its material properties. The symbol is... Indicates the first Volume of each unit, symbol Indicates the first The density of the solid material used in each unit. It can be understood that by adjusting the density of each unit... The value is found under the condition of satisfying the constraints. Minimal material distribution.

[0098] In the specific implementation of material distribution optimization based on the force transmission path distribution map, high-density areas in the force transmission path distribution map are marked as critical load-bearing areas, and low-density areas are marked as secondary load-bearing areas. In some embodiments, for the critical load-bearing areas, carbon fiber composite material grades with tensile strength and compressive strength both higher than preset thresholds are selected from a preset material database and identified as high-stiffness composite material areas. For the secondary load-bearing areas, the original lightweight foam filling structure is retained to maintain the buoyancy characteristics and low-speed aerodynamic efficiency of the body. In the critical load-bearing areas, the joint locations bearing concentrated loads are identified, and titanium alloys or aluminum alloys with high yield strength are selected from the material database as local reinforcing metal parts. Optionally, the determined high-stiffness composite material areas, lightweight foam filling structures, and local reinforcing metal parts are combined in the body structure to form a hybrid material distribution scheme.

[0099] In one embodiment of the present invention, the process of converting a hybrid material distribution scheme into detailed structural design drawings includes skin ply angles and detailed drawings of connection nodes of the internal skeleton. (See also...) Figure 3 The process involves reading the material type identifiers and spatial location information from the hybrid material distribution scheme to establish a one-to-one correspondence table between materials and structural components. Based on this table, parametric modeling of the UAV platform's skin is performed in the 3D design environment. Inputs for parametric modeling include the ply thickness and fiber orientation angle of the high-stiffness composite material. The internal skeleton of the UAV platform is then modeled as a truss structure, with locally reinforcing metal components defined as rigid connection nodes. Following parametric and truss modeling, assembly interference checks are performed to correct structural conflicts caused by variations in material thickness. The 3D model, after interference checks and corrections, is projected onto 2D engineering drawings, with details of the skin ply angles and internal skeleton connection nodes annotated to create detailed structural design drawings.

[0100] Based on the detailed structural design drawings, a manufacturing process card was compiled. The names and layup structures of all components made of high-rigidity composite materials were extracted from the detailed structural design drawings. For each component made of high-rigidity composite materials, an autoclave molding process route was developed, specifying the heating rate, holding time, and cooling rate as control parameters for the composite material's curing temperature. The dimensions of all parts composed of locally reinforced metal components were extracted from the detailed structural design drawings, and their nominal dimensions and tolerance zones were calculated. For each part composed of locally reinforced metal components, a CNC machining center was selected as the processing equipment. The calculated tolerance zones were converted into machining tolerances for the metal parts, and a machining program was compiled. The autoclave molding process route and machining program were integrated into a manufacturing process card for a heavy-load, lightweight unmanned aerial vehicle platform.

[0101] In practical implementation, a design method for a heavy-payload lightweight UAV platform involves converting a hybrid material distribution scheme into detailed structural design drawings. These drawings include details of the skin ply angles and connection nodes of the internal skeleton. The material type identifiers and spatial location information from the hybrid material distribution scheme are read to establish a one-to-one correspondence table between materials and structural components. This table records component numbers such as "fuselage keel section" and "skin on the central wing box" and their corresponding material grades such as "T800 grade carbon fiber composite material" or "TC4 titanium alloy," along with their three-dimensional spatial coordinate ranges. Based on this correspondence table, parametric modeling of the heavy-payload lightweight UAV platform's skin is performed in a three-dimensional design environment. The inputs for parametric modeling include the ply thickness and fiber orientation angles of the high-stiffness composite material. For example, the input parameters can specify a total ply thickness of 2.1 mm for a certain skin region, with a ply sequence of [0° / 45° / 90° / -45°]s. The internal skeleton of a heavy-payload lightweight UAV platform is modeled using a strut structure. In this modeling, locally reinforcing metal components are defined as rigid connection nodes. Based on the parametric and strut structure models, assembly interference checks are performed to correct structural conflicts caused by variations in material thickness, such as correcting the installation gap between the composite wing spars and the titanium alloy joints. Essentially, the 3D model, after interference checks and corrections, is projected onto a 2D engineering drawing, annotating the skin ply angles and detailed connection nodes of the internal skeleton to create a detailed structural design drawing. The ply angles are listed in a table in the corner of the drawing, and the detailed connection node drawings show the bolt hole diameter and positional tolerances.

[0102] In practice, manufacturing process cards are compiled based on detailed structural design drawings. The names and layup structures of all components made of high-rigidity composite materials are extracted from these drawings. For example, the "left outer wing upper skin" component is identified as being composed of 12 layers of unidirectional carbon fiber prepreg. For each component made of high-rigidity composite materials, an autoclave molding process route is developed, specifying the heating rate, holding time, and cooling rate as control parameters for the composite material's curing temperature. For instance, one route specifies heating to 180 degrees Celsius at a rate of 1.5 degrees Celsius per minute, holding for 120 minutes, and then cooling to below 60 degrees Celsius at a rate not exceeding 2 degrees Celsius per minute. The dimensions of all parts composed of locally reinforced metal components are extracted from the detailed structural design drawings, and their nominal dimensions and tolerance zones are calculated. For each part composed of locally reinforced metal components, a five-axis CNC machining center is selected as the machining equipment. The calculated tolerance zones are converted into machining tolerances for the metal parts, and a machining program is developed. Optionally, the autoclave molding process route and machining procedures can be integrated into a manufacturing process card for a heavy-load, lightweight UAV platform. In the manufacturing process card, the curing temperature profile of the composite material is listed graphically, and the machining tolerances of the metal parts are marked in dimensional chains. In some embodiments, the allocation of machining tolerances for the metal parts needs to consider cumulative assembly errors; a model formula for calculating statistical tolerances is as follows:

[0103] ;

[0104] Where: symbol The symbol represents the total assembly tolerance of a closed dimensional chain. Indicates the total number of components in a size chain, symbol It is the index number that makes up the ring, symbol Indicates the first The dimensional tolerances of each component ring, with symbols... Indicates the first The relative distribution coefficients of each component ring are used to describe the distribution characteristics of its machined dimensions. It can be understood that, based on this model, reasonable machining tolerances can be assigned to each locally reinforcing metal part. To ensure final assembly tolerances It meets the requirements of detailed structural design drawings. Optionally, when compiling the machining program, the cutting parameters of the CNC machining center, such as spindle speed and feed rate, need to be matched and set according to the material properties of the locally reinforcing metal parts. In some embodiments, the final integrated manufacturing process card is a structured document that clearly specifies the inputs, outputs, equipment parameters, and quality checkpoints for each manufacturing step.

[0105] In one embodiment of the present invention, after constructing an initial overall layout scheme based on mission profile parameters, the initial overall layout scheme is imported into an aerodynamic disturbance analysis module to simulate the disturbance of the wing flow field after the payload is mounted. Pressure coefficient distribution data output by the aerodynamic disturbance analysis module is collected to identify high-risk areas of airflow separation on the wing surface. Based on the location of these high-risk areas, the installation position and deflection angle of the wing trailing edge flaps are adjusted in the initial overall layout scheme. The adjusted parameters of the wing trailing edge flaps are fed back to the initial overall layout scheme to update its geometric configuration data. When updating the geometric configuration data, the adjustment parameters of the wing trailing edge flaps are read, including the deflection angle value and the offset distance relative to the wing chord line. In the geometric model of the initial overall layout scheme, the coordinates of the drive hinge point of the wing trailing edge flaps are located. Based on the deflection angle value and the offset distance, the new coordinates of the trailing edge flap tip outline point in the airframe coordinate system are calculated. The new tip outline point coordinates replace the old ones in the geometric model, reconstructing the three-dimensional solid model of the wing trailing edge flaps. The original wing trailing edge flap model in the initial overall layout scheme was replaced with the reconstructed 3D solid model of the wing trailing edge flap, thus completing the update of the geometric configuration data.

[0106] See Figure 4 In the aerodynamic interference analysis of flaps in a lightweight UAV with heavy load, the variation of the wing surface pressure coefficient with the flap deflection angle intuitively reflects the evolution of airflow separation risk. The curve uses the flap deflection angle as the independent variable and the pressure coefficient at the wing surface measurement point as the dependent variable, with a pressure coefficient of -0.8 as the threshold for judging airflow separation risk. When the flap deflection angle is -5° downward, the wing surface pressure coefficient is as low as -1.2, the negative pressure peak is obvious, and the airflow adhesion stability is poor. As the flap deflection angle gradually increases in the positive direction, the pressure coefficient increases accordingly, the absolute value of the negative pressure continues to decrease, and the wing surface load distribution... When the deflection angle reaches 5°, the pressure coefficient touches the separation risk threshold, and the airflow separation risk enters a critical state. After exceeding 10°, the pressure coefficient further increases to -0.6, the negative pressure level is significantly reduced, the airflow separation risk is effectively suppressed, and the aerodynamic load distribution on the wing surface is more reasonable. This curve provides a direct quantitative basis for the optimized design of the flap deflection angle. By controlling the flap deflection angle to be above 5°, the wing surface pressure coefficient can be maintained above the separation risk threshold, effectively avoiding airflow separation and ensuring the aerodynamic stability and control safety of the UAV under heavy load conditions.

[0107] In practice, the aerodynamic interference analysis module employs a Reynolds-averaged Navier-Stokes equation solver to calculate the overall airflow field within a typical flight speed range and under configurations with external loads. It collects pressure coefficient distribution data output by the module and identifies high-risk areas for airflow separation on the wing surface. This identification process involves analyzing the slope and pressure recovery region of the pressure coefficient curve on the upper wing surface. Based on the location of these high-risk areas, the installation position and deflection angle of the wing trailing edge flaps are adjusted in the initial overall layout scheme. The adjusted trailing edge flap parameters are then fed back to the initial overall layout scheme to update its geometric configuration data. Table 1 shows pressure coefficient distribution segments at different airfoil positions on the wing surface before flap adjustment, used to assist in identifying the initial region of airflow separation.

[0108] Table 1: Distribution of pressure coefficients at different airfoil positions on the wing surface before flap adjustment (segment table)

[0109] 0.10 -1.25 0.15 0.30 -2.80 0.10 0.55 -1.90 -0.20 0.70 -0.60 -0.10 0.85 0.10 0.05

[0110] In some embodiments, the process of updating the geometric configuration data begins by reading the adjustment parameters of the wing trailing edge flap, including the deflection angle value and the offset distance relative to the wing chord. In the geometric model of the initial overall layout scheme, the coordinates of the drive hinge point of the wing trailing edge flap are located; the drive hinge point coordinates are defined as a fixed point in the body coordinate system. Based on the deflection angle value and the offset distance, the new coordinates of the end profile point of the wing trailing edge flap in the body coordinate system are recalculated, involving a rigid coordinate transformation about the hinge point. One formula for calculating the new coordinates of the end profile point is expressed as:

[0111] ;

[0112] Where: symbol This represents the new coordinate components of the trailing edge flap tip profile point in the airframe coordinate system, denoted by the symbol. Represents the original coordinate components of the end contour point, with the symbol... Represents the coordinates of the driving hinge point, symbol Represents the components of the offset distance relative to the wing chord in the three coordinate directions, with the symbol... Represents a rotation about an axis defined by the driving hinge point. The three-dimensional rotation matrix of the angle, symbol This represents the deflection angle value of the wing trailing edge flap. Optionally, the old trailing edge profile point coordinates in the geometric model are replaced with the new ones to reconstruct the 3D solid model of the wing trailing edge flap. The reconstruction process is completed using the surface lofting and stitching functions of CAD software. In some embodiments, the reconstructed 3D solid model of the wing trailing edge flap replaces the original wing trailing edge flap model in the initial overall layout scheme, completing the update of the geometric configuration data. The updated geometric configuration data will be used as input for subsequent mass characteristic calculations. It can be understood that updating the geometric configuration data is a step in the iterative design process, used to respond to aerodynamic analysis results and correct the initial design. Optionally, offset distance. The values ​​are determined based on the location and size of high-risk areas of airflow separation in order to optimize the control effect of flaps on wing circulation.

[0113] In one embodiment of the present invention, before converting the hybrid material distribution scheme into detailed structural design drawings, buckling stability is checked on the high-stiffness composite material region in the hybrid material distribution scheme. The load data used for buckling stability check is the empty weight of the entire aircraft in the initial mass characteristic dataset. If the check result does not meet the preset stability margin requirement, the number of local ply layers in the high-stiffness composite material region is increased, and the buckling stability check is performed again. When the check result meets the stability margin requirement, the number of ply layers in the high-stiffness composite material region at this time is confirmed as the final design value. The high-stiffness composite material region data containing the final design value is substituted into the hybrid material distribution scheme, and then the step of converting it into detailed structural design drawings is executed. After compiling the manufacturing process card, the composite material curing temperature curve and the machining tolerance range of the metal parts are extracted from the manufacturing process card to form quantifiable quality control indicators. The quality control indicators are entered into the production quality control system of the UAV platform as the inspection benchmark for subsequent trial production batches. During the production process, the curing temperature monitoring data and machining dimension detection data of the actual trial production parts are collected and compared with the entered quality control indicators. The deviation values ​​in the comparison results are fed back to the structural topology optimization model as input data for adjusting constraints in the next design iteration.

[0114] See Figure 5In the production quality control process of the heavy-load lightweight UAV platform, the changing trends of curing temperature deviation and processing tolerance deviation with the trial production batches intuitively reflect the convergence effect of process iterative optimization. In the figure, the blue curve represents the curing temperature deviation of composite materials, and the orange curve represents the processing tolerance deviation of metal parts. Both curves show a monotonically decreasing trend with the increase of trial production batches. In the first batch of trial production, the curing temperature deviation was 5℃ and the processing tolerance deviation was 0.03mm, which was a relatively high level of process deviation. After deviation feedback and process parameter iterative adjustment by the production quality control system, the curing temperature deviation of the second batch was reduced to 3℃ and the processing tolerance deviation was reduced to 0.02mm. The third batch was further optimized to 1℃ and 0.015mm. By the fourth batch, the curing temperature deviation had converged to 0.5℃ and the processing tolerance deviation had converged to 0.01mm, all of which met the preset quality control index requirements. This curve verifies the effectiveness of the closed-loop control mechanism of feeding back the actual trial production deviation to the structural topology optimization model to iteratively adjust the constraints, and provides quantitative support for the continuous improvement of the UAV platform manufacturing process.

[0115] In practice, buckling stability verification is performed through finite element eigenvalue buckling analysis. The applied load condition is the ultimate flight load calculated based on the empty weight of the entire aircraft. If the verification result does not meet the preset stability margin requirement, the number of local ply counts in the high-stiffness composite material region is increased, and the buckling stability verification is performed again. When the verification result meets the stability margin requirement, the number of ply counts in the high-stiffness composite material region at this point is confirmed as the final design value. The high-stiffness composite material region data containing the final design value is substituted into the hybrid material distribution scheme, and then the step of converting it into detailed structural design drawings is performed. The stability margin requirement is defined by a dimensionless stability safety factor, the calculation formula of which is:

[0116] ;

[0117] Where: symbol The calculated stability safety factor is represented by the symbol. The symbol represents the critical buckling load of the structure calculated through eigenvalue buckling analysis. This represents the design ultimate load acting on the main load-bearing structure, determined based on the initial mass characteristic data set, including the aircraft's empty weight and flight envelope. It can be understood that the preset stability margin requirement is equivalent to the specified stability safety factor. It must be greater than a minimum threshold If not satisfied If the condition is not met, the verification result is deemed not to meet the preset stability margin requirement. In some embodiments, the number of local plies in the high-stiffness composite material region is increased. Specifically, in the finite element model, for the maximum displacement region corresponding to the identified low-order buckling mode, a ply in a specific direction is added in the ply definition.

[0118] After compiling the manufacturing process card, the curing temperature curve of the composite material and the machining tolerance range of the metal parts are extracted from the manufacturing process card to form quantifiable quality control indicators. In specific implementation, the curing temperature curve of the composite material is quantified into the heating rate value, the target temperature value and duration of the holding stage, and the cooling rate value. The machining tolerance range of the metal parts is quantified into the upper and lower deviations of the dimensions. The quality control indicators are entered into the production quality control system of the heavy-load lightweight UAV platform as the inspection benchmark for subsequent trial production batches. During the production process, the curing temperature monitoring data and machining dimension detection data of the actual trial-produced parts are collected and compared with the entered quality control indicators. Optionally, the collected actual curing temperature monitoring data comes from the real-time readings of multiple thermocouples in the autoclave, and the collected actual machining dimension detection data comes from the inspection report of the coordinate measuring machine. In some embodiments, the deviation value in the comparison result is calculated as the difference between the measured data and the target value in the quality control indicator. The deviation value in the comparison result is fed back to the structural topology optimization model as input data for adjusting the constraints in the next design iteration. It is understandable that the feedback process involves converting deviation values ​​into parameter corrections that the model can recognize. For example, if the actual holding time for the composite material's curing temperature is consistently lower than the range specified in the process card, the input value of the allowable stress constraint for the material will be adjusted accordingly in the structural topology optimization model. Optionally, after receiving the feedback deviation values, the structural topology optimization model will recalculate the force transmission path distribution map based on updated material properties or manufacturing precision constraints in subsequent optimization loops.

[0119] The above are merely preferred embodiments of the present invention and are not intended to limit the present invention in any other way. Any person skilled in the art may make changes or modifications to the above-disclosed technical content to create equivalent embodiments that can be applied to other fields. However, any simple modifications, equivalent changes, and modifications made to the above embodiments based on the technical essence of the present invention without departing from the scope of the present invention shall still fall within the protection scope of the present invention.

Claims

1. A method for designing a large load lightweight unmanned aerial vehicle platform, characterized in that, The method includes: Obtain the mission profile parameters of the large-payload lightweight unmanned aerial vehicle platform, including the maximum takeoff weight, payload mass, endurance index, and typical flight speed range; Based on the mission profile parameters, an initial overall layout scheme for the heavy-payload lightweight UAV platform is constructed. The initial overall layout scheme is subjected to mass characteristic calculation to obtain an initial mass characteristic dataset containing the center of gravity coordinates, rotational inertia matrix and the empty weight of the whole machine; The initial mass characteristic dataset is input into the structural topology optimization model. With the goal of minimizing the empty weight of the entire aircraft and the constraints of the bending moment generated by the payload mass and the axial force generated by the maximum takeoff weight, a force transmission path distribution diagram of the main load-bearing structure is generated. Based on the force transmission path distribution diagram, the material distribution of the main load-bearing structure is optimized to obtain a hybrid material distribution scheme consisting of a high-stiffness composite material region and locally reinforced metal parts; The hybrid material distribution scheme is converted into detailed structural design drawings of the heavy-load lightweight UAV platform. Based on the detailed structural design drawings, a manufacturing process card for the heavy-load lightweight UAV platform is prepared. The manufacturing process card specifies the curing temperature of the composite material and the machining tolerances of the metal parts.

2. The design method for a heavy-payload lightweight unmanned aerial vehicle platform according to claim 1, characterized in that, The initial overall layout scheme is subjected to mass characteristic calculation to obtain an initial mass characteristic dataset containing the center of gravity coordinates, rotational inertia matrix, and overall empty weight, including: The initial overall layout scheme defines the fuselage aspect ratio, wing aspect ratio, and power system installation location; In the initial overall layout scheme, geometric modeling is performed on each component of the airframe to generate a geometric model that includes the fuselage outer curved surface, the wing planar shape, and the landing gear strut positions; Assign corresponding material density properties to each part of the geometric model, wherein the material density properties are derived from a preset material database; By discretizing the geometric model, the centroid position and volume of each discrete unit are calculated, and the centroid position and volume data of all discrete units are summarized. Based on the centroid position and volume data of all discrete units, the centroid coordinates and empty weight of the entire machine of the initial overall layout scheme are calculated. Based on the mass distribution of the geometric model and the spatial position of the discrete units, the rotational inertia components around the three main axes of the body are calculated and assembled to form the rotational inertia matrix. The center of gravity coordinates, moment of inertia matrix, and empty weight of the entire machine are packaged and encapsulated to form the initial mass characteristic dataset.

3. The design method for a lightweight unmanned aerial vehicle platform with large payload according to claim 2, characterized in that, The initial mass characteristic dataset is input into the structural topology optimization model. With the goal of minimizing the empty weight of the entire aircraft and constrained by the bending moment generated by the payload mass and the axial force generated by the maximum takeoff weight, a force transmission path distribution diagram of the main load-bearing structure is generated, including: The empty weight of the aircraft is extracted from the initial mass characteristic dataset as the benchmark value for optimization, and the bending moment generated by the payload mass and the axial force generated by the maximum takeoff weight are extracted as the working condition load boundary. The working condition load boundary and the benchmark value are imported into the structural topology optimization model, and the number of optimization iterations is set to a preset fixed value; In the structural topology optimization model, the empty weight of the whole machine is taken as the objective function, and the structural stiffness under the working condition load boundary is taken as the constraint condition to start the iterative calculation process; In each iterative calculation, the material's state of existence in space is adjusted based on the current stress distribution cloud map and strain energy density distribution cloud map; When the number of iterations reaches a preset fixed value, the iteration process is terminated, and the boundary contours of the high-density and low-density regions of the material at this time are extracted. Map the boundary contour back to the body coordinate system of the heavy-load lightweight UAV platform and draw the force transmission path distribution diagram of the main load-bearing structure; The steps for constructing the structural topology optimization model include: Obtain a three-dimensional geometric model of the initial overall layout scheme of the large-payload lightweight UAV platform, wherein the three-dimensional geometric model includes the original design space of the main load-bearing structure; Material properties are applied to the three-dimensional geometric model, and the entire design space is discretized into a uniformly distributed finite element mesh; Multiple load application points and displacement constraint points are defined on the boundary of the design space. The load application points correspond to the working load boundaries in the initial mass characteristic dataset, and the displacement constraint points correspond to the fixing methods of the landing gear connection points and the power system mounting points. The objective function of the structural topology optimization model is set as minimizing the overall empty weight of the machine; Multiple optimization constraints are input into the structural topology optimization model. The optimization constraints include that the maximum structural deformation under the bending moment generated by the effective payload mass does not exceed a preset value, and the maximum structural stress under the axial force generated by the maximum takeoff weight does not exceed the allowable stress of the material. Define the iterative convergence criterion for the structural topology optimization model, wherein the iterative convergence criterion is that the change in the material volume fraction of the design space calculated in two adjacent iterations is less than a preset threshold.

4. The design method for a lightweight unmanned aerial vehicle platform with a large payload according to claim 3, characterized in that, Based on the force transmission path distribution diagram, the material distribution of the main load-bearing structure is optimized to obtain a hybrid material distribution scheme consisting of a high-stiffness composite material region and locally reinforced metal parts, including: The high-density areas in the force transmission path distribution map are marked as critical load-bearing areas, and the low-density areas are marked as secondary load-bearing areas. For the critical load-bearing area, carbon fiber composite material grades with tensile strength and compressive strength both higher than the threshold are selected from the preset material database and identified as the high stiffness composite material region; For the secondary load-bearing area, the original lightweight foam filling structure is retained to maintain the buoyancy characteristics and low-speed aerodynamic efficiency of the body; In the critical load-bearing area, the joint locations bearing concentrated loads are identified, and titanium alloys or aluminum alloys with high yield strength are selected from the material database as the local reinforcing metal parts. The determined positions of the high-rigidity composite material region, the lightweight foam filling structure, and the locally reinforcing metal parts within the body structure are combined to form the hybrid material distribution scheme.

5. The design method for a heavy-payload lightweight unmanned aerial vehicle platform according to claim 4, characterized in that, The hybrid material distribution scheme is transformed into detailed structural design drawings for the heavy-load lightweight UAV platform, including: The detailed structural design drawings include skin ply angles and detailed drawings of the connection nodes of the internal skeleton; Read the material type identifiers and spatial location information in the hybrid material distribution scheme, and establish a one-to-one correspondence table between materials and structural components; Based on the one-to-one correspondence table, the skin of the heavy-load lightweight UAV platform is parametrically modeled in the three-dimensional design environment. The inputs for the parametric modeling include the layup thickness and fiber orientation angle of the high-stiffness composite material. The internal skeleton of the heavy-load lightweight UAV platform is modeled as a rod structure, and in the rod structure modeling, the local reinforcing metal parts are defined as rigid connection nodes; Based on the parametric modeling and the rod structure modeling, an assembly interference check is performed to correct structural conflicts caused by changes in material thickness. The 3D model, after interference checking and correction, is projected onto a 2D engineering drawing, and the skin ply angle and the connection node details of the internal skeleton are marked to form the detailed structural design drawings.

6. The design method for a heavy-payload lightweight unmanned aerial vehicle platform according to claim 5, characterized in that, Based on the detailed structural design drawings, a manufacturing process card for the heavy-payload lightweight UAV platform is prepared, including: The names of all components made of the high-rigidity composite material and their ply structures are extracted from the detailed structural design drawings. For each component made of the high-rigidity composite material, an autoclave molding process route is formulated, in which the heating rate, holding time and cooling rate are set as control parameters for the curing temperature of the composite material. The dimensions of all parts composed of the locally reinforcing metal parts are extracted from the detailed structural design drawings, and their nominal dimensions and tolerance zones are calculated. For each part composed of the aforementioned locally reinforcing metal parts, a CNC machining center is selected as the machining equipment. The calculated tolerance zone is converted into the machining tolerance of the metal parts, and a machining program is compiled. The autoclave forming process route and the processing procedure are integrated into a manufacturing process card for the heavy-load lightweight UAV platform.

7. The design method for a lightweight unmanned aerial vehicle platform with large payload according to claim 6, characterized in that, After constructing the initial overall layout scheme of the heavy-payload lightweight UAV platform based on the mission profile parameters, the following steps are also included: The initial overall layout scheme is imported into the aerodynamic interference analysis module to simulate the disturbance of the wing flow field after the payload is mounted. The pressure coefficient distribution data output by the aerodynamic interference analysis module is collected to identify high-risk areas of airflow separation on the wing surface; Based on the location of the high-risk area of ​​airflow separation, the installation position and deflection angle of the wing trailing edge flaps are adjusted in the initial overall layout scheme; The adjusted parameters of the wing trailing edge flaps are fed back to the initial overall layout scheme, updating the geometric configuration data of the initial overall layout scheme.

8. The design method for a large-payload lightweight unmanned aerial vehicle platform according to claim 7, characterized in that, The adjusted parameters of the wing trailing edge flaps are fed back to the initial overall layout scheme, updating the geometric configuration data of the initial overall layout scheme, including: Read the adjustment parameters of the trailing edge flap of the wing, including the deflection angle value and the offset distance relative to the wing chord line; In the geometric model of the initial overall layout scheme, locate the coordinates of the drive hinge point of the trailing edge flap of the wing; Based on the deflection angle and the offset distance, the new coordinates of the end profile point of the wing trailing edge flap in the body coordinate system are recalculated. The three-dimensional solid model of the wing trailing edge flap is reconstructed by replacing the old trailing edge point coordinates in the geometric model with the new trailing edge point coordinates. The reconstructed 3D solid model of the wing trailing edge flap replaces the original wing trailing edge flap model in the initial overall layout scheme, thus completing the update of the geometric configuration data.

9. The design method for a lightweight unmanned aerial vehicle platform with large payload according to claim 8, characterized in that, Before converting the hybrid material distribution scheme into detailed structural design drawings for the heavy-load lightweight UAV platform, the following steps are also included: Buckling stability is checked in the high-stiffness composite material region of the hybrid material distribution scheme, and the load data used for buckling stability check is the empty weight of the whole machine in the initial mass characteristic dataset; If the verification result does not meet the preset stability margin requirement, the number of local layups in the high stiffness composite material region is increased, and the buckling stability verification is performed again. When the verification result meets the stability margin requirement, the number of layups in the high-stiffness composite material region at this time is confirmed as the final design value. Substitute the high-stiffness composite material region data containing the final design value into the hybrid material distribution scheme, and then perform the step of converting it into detailed structural design drawings.

10. The design method for a heavy-payload lightweight unmanned aerial vehicle platform according to claim 9, characterized in that, After compiling the manufacturing process card for the aforementioned heavy-payload lightweight unmanned aerial vehicle platform, the following steps are also included: The curing temperature curve of the composite material and the machining tolerance range of the metal part are extracted from the manufacturing process card to form quantifiable quality control indicators. The quality control indicators are entered into the production quality control system of the heavy-payload lightweight UAV platform as the inspection benchmark for subsequent trial production batches. During the production process, the curing temperature monitoring data and machining dimension detection data of the actual prototype parts are collected and compared with the recorded quality control indicators; The deviation values ​​in the comparison results are fed back to the structural topology optimization model as input data for adjusting constraints in the next design iteration.

Images