Lightweight optimization method for wing with high aspect ratio

Abstract

The invention discloses a lightweight optimization method for a high-aspect-ratio wing, and relates to the fields of structural dynamics, aircraft structural design and the like. Structural dynamics and flutter characteristics are comprehensively considered, and the problems that in the lightweight design process of the flexible wing, the dynamic characteristic adjustment space is small, and optimization methods are few are solved. The method comprises the following steps: 1, establishing a high-fidelity structure finite element model of a flexible wing model; 2, establishing a size optimization model of the flexible wing; 3, solving the optimization model based on finite element software and the finite element model; and 4, modifying the wing geometric model according to the optimized design variable value. The optimization model established in the design process takes structural dynamics and flutter characteristics as constraints, so that the optimization result better conforms to the characteristics of the flying wing layout aircraft, and the structural dynamics characteristics of the aircraft can be further improved.

Description

Technical Field

[0001] The present invention relates to the fields of structural dynamics, aircraft structural design, and in particular to a flexible wing lightweight structural dynamics design technology. Background Art

[0002] In recent years, with the aviation industry's demand for advanced aircraft in terms of aerodynamics, stealth and other performance, flying wing layout aircraft have received widespread attention. This type of aircraft generally adopts a flexible wing structure with a large aspect ratio, which has good lift-drag characteristics; at the same time, it is manufactured using lightweight materials, which effectively reduces the weight of the structure and can reduce fuel consumption. However, due to the light weight and high flexibility of this wing, the wing will produce large bending and torsional deformations under the action of flight loads, which can easily cause complex structural dynamics problems, such as flutter. Therefore, when designing a flexible wing structure, the influence of structural dynamic characteristics must be taken into account, which brings new challenges to the lightweight design of flexible wings.

[0003] At present, when the academic and industrial circles are designing the lightweight internal structure of flexible wing models, they usually adopt a skeleton structure layout including wing beams, longitudinal walls, stiffeners, wing ribs, etc. The shapes and distribution positions of these components are generally fixed, so there is little space for lightweight design, and it is not convenient to adjust the dynamic characteristics of the aircraft.

[0004] Researchers and technicians often use optimization methods when designing lightweight flexible wings. The optimization method commonly used in existing research is to first perform aerodynamic optimization on the wing, then optimize the size with the goal of weight reduction, taking aerodynamic loads into consideration, and taking structural strength as a constraint to obtain an optimization result with good aerodynamic and strength characteristics. This optimization method pays less attention to the structural dynamic characteristics and easily ignores the impact of flutter on flexible wings. In addition, the finite element model of the wing structure on which the existing optimization work is based has generally been simplified to a certain extent, so that the optimization results may deviate from the actual model.

[0005] In addition, with the continuous development of materials and manufacturing technology, additive manufacturing technology has become one of the important means of future aircraft design and manufacturing due to its advantages such as simple processing procedures and fast molding. For flying wing layout aircraft, the use of additive manufacturing can accurately control the shape and position of the internal force-bearing structure of the aircraft, which facilitates the lightweight design of the structure and the adjustment of the mechanical properties of the model.

[0006] Therefore, it is of great significance for the lightweight design of flexible wings to design an internal structure of a flexible wing that is easy to optimize and convenient to adjust the dynamic characteristics, and to establish a flexible wing lightweight optimization method that considers the structural dynamics and flutter characteristics. Summary of the invention

[0007] In view of the above problems, the present invention proposes a lightweight optimization method for a large aspect ratio wing, which comprehensively considers the structural dynamics and flutter characteristics, and solves the problems of small adjustment space for dynamic characteristics and few optimization methods in the lightweight design process of flexible wings.

[0008] The technical solution of the present invention is: comprising the following steps:

[0009] Step 1, establishing a high-fidelity structural finite element model of the flexible wing model;

[0010] First, the designed high-fidelity wing geometry model is imported into the finite element software. Then, the finite element mesh of the model is accurately established according to the shape and size of the geometry model, and the mesh unit attributes and material attributes are assigned.

[0011] Step 2, establishing a size optimization model of the flexible wing;

[0012] The objective function of the optimization model is the first n orders of the flexible wing, where n is a positive integer. The sum of the squares of the deviations between the natural frequencies of the modes other than the in-plane modes and the scaled natural frequencies of the prototype wing is required to be minimum, as shown in the following formula:

[0013] where i is the order, ω αi is the natural frequency of the i-th order mode obtained experimentally, that is, the natural frequency after optimization, ω mi is the natural frequency of the i-th order mode obtained by finite element, that is, the natural frequency of the prototype wing;

[0014] The design variables of the optimization model include the thickness of each unit cell filled inside the wing, the thickness of the winglet, the thickness of the skin, and the thickness of the spar; the constraints of the optimization model include stress constraints, natural frequency constraints, vibration mode constraints, and flutter speed constraints;

[0015] The upper and lower boundaries of the design variables and constraints are given. The upper and lower boundaries of the design variables are given according to the model processing technology limitations and experience. Among the upper and lower boundaries of the constraints, the stress constraint is given according to the strength of the model material; the natural frequency constraint is given according to the natural frequency of the physical aircraft corresponding to the flexible wing model after scaling calculation; the vibration mode constraint is given according to the vibration mode coordinates of the physical aircraft corresponding to the flexible wing model after scaling calculation; the flutter speed constraint is given according to the expected flight conditions;

[0016] Step 3: Solve the optimization model based on finite element software and finite element model;

[0017] Use the optimization module of the finite element modeling software Patran to input the optimization model established in step 2 into it; then submit the finite element model to the finite element analysis software MSC.Nastran, perform optimization calculation according to the upper and lower boundaries of the given design variables and constraints in step 2, and optimize the design variables;

[0018] If the solution does not converge, reselect the upper and lower boundaries of the design variables and constraints and submit the calculation again; after the solution converges, derive the design variable values ​​obtained from the last iteration;

[0019] Step 4: modify the wing geometry model according to the optimized design variable values;

[0020] According to the optimized natural frequency, the deviation of the natural frequency is calculated, and the weight and length of the wing are modified so that the values ​​of the corresponding dimensions of the geometric model are equal to the optimized values.

[0021] The interior of the wing is filled with a lattice structure composed of unit cells, the lattice structure is covered with a skin, and the unit cell is formed by connecting the bottom surfaces of two hexagonal pyramid-shaped frames.

[0022] The wing comprises a wing-body fusion part 1, a wing section 2 and a winglet 4 which are connected in sequence. One side of the wing-body fusion part 1 is used to connect to the fuselage of the aircraft, and the other side is fixedly connected to the winglet 4 through two wing sections 2. Two control surfaces 3 are respectively installed on each wing section 2, and a steering gear for controlling the rotation of the control surface 3 is installed inside the wing section 2.

[0023] The wing section 2 and the control surface 3 are processed by a 3D printer, and the interior thereof is filled with unit cells to form a lattice structure, and a skin is coated outside the lattice structure;

[0024] The wing-body fusion portion 1 and the wing section 2, adjacent wing sections 2, and the wing section 2 and the winglet 4 are all connected via a separation surface connection structure.

[0025] Two adjacent structural members are connected by the separation surface connection structure, and the separation surface connection structure is as follows: a slot 12 along the chord direction of the wing is opened on one of the structural members, and a ridge 13 along the chord direction of the wing is provided on the other structural member. After the ridge is assembled in the slot along the chord direction, it is locked by a vertically arranged fixing screw passing through the ridge 13.

[0026] Compared with the prior art, the present invention has the following beneficial effects:

[0027] 1. The interior of the flexible wing is filled with a lattice structure composed of unit cells. The unit cells are hollow thin-walled structures that can effectively reduce the weight of the wing. If the thickness of each unit cell is designed individually, the mass and stiffness distribution of the wing can be changed, thereby allowing a larger adjustment space for the aircraft's dynamic characteristics.

[0028] Second, when the model is processed by additive manufacturing, by adjusting the angle between the side of the unit cell and the plane perpendicular to the bottom surface, no support material will be generated in the unit cell when printing along the chord direction of the wing, thereby facilitating the overall printing of the wing and reducing the processing workload.

[0029] 3. Use high-fidelity finite element models for analysis, and for additive manufacturing materials, use anisotropic models that are more in line with the actual characteristics of the materials to define material properties, which can reduce the error between the simulation results and the physical model performance.

[0030] 4. The optimization model established during the design process takes into account structural dynamics and flutter characteristics as constraints, making the optimization results more consistent with the characteristics of flying wing layout aircraft, and also facilitating further improvement of the structural dynamic characteristics of the aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0031] Figure 1 It is a top view of the structural layout of the high aspect ratio wing model;

[0032] Figure 2 This is a bottom view of the high aspect ratio wing model structure layout;

[0033] Figure 3 This is a schematic diagram of the dot matrix structure used inside the wing;

[0034] Figure 4 It is the top view and side view of the control surface shape;

[0035] Figure 5 It is a schematic diagram of the assembly of the control surface and the wing section connection structure;

[0036] Figure 6a This is a schematic diagram of the structure of the two halves of the wingtip winglet. Figure 1 , Figure 6b This is a schematic diagram of the structure of the two halves of the wingtip winglet. Figure 2 ;

[0037] Figure 7a This is a schematic diagram of the structure on one side of the separation surface connection structure Figure 1 , Figure 7b This is the structural diagram of this side. Figure 2 ;

[0038] Figure 8a This is a schematic diagram of the structure on the other side of the separation surface connection structure Figure 1, Figure 8b This is the structural diagram of this side. Figure 2 ;

[0039] Fig. 9 It is a schematic diagram of a circular shaft with a spring used in the connection mechanism between the control surface and the wing section;

[0040] Fig.10 It is a schematic diagram of the appearance of the unit cell structure provided by the present invention;

[0041] Fig.11 is based on Fig.10 Schematic diagram of the flexible wing structure obtained from the mesocell unit;

[0042] Fig.12 It is a flow chart of a method for designing the lightweight structure dynamics of a flexible wing;

[0043] Fig.13 It is the vibration shape of the first five modes of the optimized model except the in-plane mode;

[0044] Fig.14 It is the scalar function and frequency response curve at the origin under each working condition;

[0045] Fig.15 is the modal formation diagram of each order without winglets after mass normalization;

[0046] In the figure, 1 is the wing-body fusion part, 2 is the wing section, 3 is the control surface, 4 is the wingtip winglet, 5 is the servo cabin cover, 6 is the hinge hole, 7 is the limit hole, 8 is the threaded hole, 9 is the mounting hole, 10 is the bearing, 11 is the reinforcement rib, 12 is the groove, 13 is the convex ridge, 17 is the aluminum alloy round shaft, and 18 is the spring. DETAILED DESCRIPTION

[0047] In order to clearly illustrate the technical features of this patent, this patent is elaborated in detail below through a specific implementation method and in combination with its accompanying drawings.

[0048] In order to make the additively manufactured high aspect ratio wing easy to connect and assemble, the installation of various parts on the high aspect ratio wing can be completed accurately and efficiently. Figure 1 and Figure 2 As shown, the model includes a wing-body fusion part 1, two wing sections 2, four control surfaces 3 and a winglet 4. The span of the whole model is 1.6m.

[0049] One side of the wing-body fusion part 1 is used to connect to the fuselage of the aircraft, and the other side is fixedly connected to the wingtip winglet 4 through two wing sections 2. Two control surfaces 3 are respectively installed on each wing section 2, and a servo for controlling the rotation of the control surface 3 is installed inside the wing section 2.

[0050] Among them, the wing uses Figure 3 The lattice structure composed of a unit cell is filled, the wing skin is placed at the lattice, and the part where the lattice and the skin intersect is cut to obtain the internal filling structure of the wing. The wing section 2 and the control surface 3 are made by additive manufacturing process, and after forming, they include unit cells and skin.

[0051] Each wing section 2 has two steering gear compartments in the middle of the lower surface. Figure 2 As shown, the interior is used to install a steering gear for controlling the deflection of the control surface, and a steering gear cabin cover 5 for sealing is fixedly installed outside the cabin.

[0052] like Figure 5 As shown, the trailing edge of the wing section 2 is provided with a receiving groove for accommodating the control surface 3, and the inner wall of the receiving groove is provided with a mounting hole 9 along the span direction, the diameter of the mounting hole is 10 mm, and each hole is provided with a rolling bearing 10 for rotationally connecting with the shaft head of the control surface 3.

[0053] The shape of the control surface 3 is as follows: Figure 4 As shown. The airfoil of the control surface 3 is obtained by dividing the airfoil into two parts near the trailing edge of the wing airfoil and taking the part including the trailing edge. The front part of the control surface 3 has a hinge hole 6 along the span direction, with a hole diameter of 6.8mm; after the rotating shaft is installed in the hinge hole 6 and the mounting hole 9, the control surface 3 is rotatably connected to the receiving groove of the wing section 2.

[0054] There is a tapered limit hole 7 at each end of the control surface 3, which points to the rotating shaft. After the limit pin is installed in the limit hole 7, the control surface 3 and the rotating shaft are relatively fixed; three threaded holes 8 distributed in a triangular shape are also opened in the middle of the control surface 3, which are used to install the transmission mechanism of the steering gear so as to accept the control of the steering gear to flip up and down.

[0055] The wingtip winglet 4 is divided into two half winglets in the vertical direction from the chord. Figure 6a , 6b As shown, bolt holes are arranged around the side edges of the two half winglets for connecting the two half winglets. The interior of the two half winglets is provided with reinforcing ribs 11, as shown in FIG. Figure 6a As shown, a transition section for connecting the wing sections is provided on the outer wall of one of the half winglets.

[0056] In addition, this embodiment also includes a connection structure between a control surface and a wing section and an assembly diagram thereof as shown in FIG. Figure 5 The rotating shaft comprises two identical aluminum alloy round shafts 17 and a spring 18, as shown in FIG. Fig. 9 The diameter of the round shaft 17 is 6 mm, and a long slot for connecting a limit pin is provided on the aluminum alloy round shaft 17 , and the spring 18 is fixedly connected between the two aluminum alloy round shafts 17 .

[0057] When installing the control surface 3, first install the two aluminum alloy round shafts 17 equipped with springs 18 into the hinge holes 6 along the span direction of the control surface 3, and press the two ends of the aluminum alloy round shafts 17 by hand so that they are completely located inside the through holes; then move the control surface 3 so that the axis of the hinge hole 6 is aligned with the axis of the hinge hole 6. Figure 5 The axes of the mounting holes 9 in the control surface 3 coincide with each other, and then release the grip, so that the two aluminum alloy round shafts 17 extend into the mounting holes 9 at both ends of the wing section mounting part along the span direction under the action of the restoring force of the springs 18; finally, limit pins are placed in the limit holes 7 at both ends of the surface of the control surface 3, and the pins are inserted into the long grooves on the surface of the aluminum alloy round shafts 17 at the same time, and the installation of the control surface can be completed.

[0058] In this case, adjacent wing sections 2 are connected via a separation surface connection structure, and the connection structure and assembly effect diagram thereof are shown in FIGS. 7 and 8 .

[0059] The separation surface connection structure is as follows: a slot 12 is provided along the chord direction on one of the wing segments 2, and a ridge 13 is provided along the chord direction on the other wing segment 2. After the ridge is assembled in the slot along the chord direction, it is locked by a vertically arranged fixing screw passing through the ridge 13.

[0060] The top and bottom surfaces of the ridge 13 are parallel planes; the ends of the slot 12 and the ridge 13 are in a broken line shape, and the surface of the separation surface connection structure is arc-shaped after the two are plugged in.

[0061] In this way, the ridges 13 and slots 12 arranged in the chord direction are assembled and then locked by the vertical fixing screws. Finally, the structural members on both sides of the separation surface can be fully positioned in the chord direction, span direction and vertical direction through the separation surface connection structure, which is not only simple to assemble, but also accurate in positioning and can be repeatedly disassembled.

[0062] The wing-body fusion part 1 and the wing section 2 are also connected by the above-mentioned separation surface connection structure. Specifically, a slot 12 can be set on the wing-body fusion part 1, and a ridge 13 can be set on the adjacent wing section 2, and finally locked by fixing screws.

[0063] The wing section 2 and the winglet 4 are also connected by the separation surface connection structure. Specifically, a slot 12 can be provided on the wing section 2, and a ridge 13 can be provided on the winglet 4, and finally locked by fixing screws.

[0064] The separation surface connection structure mentioned above can reduce the number of fasteners used in connecting the wing sections, making the installation of the wing sections relatively simple; and the wing sections connected using this method can be disassembled repeatedly. The connection structure between the control surface and the wing section provided by the present invention is relatively simple to assemble, requires fewer separation surfaces, and can be realized through additive manufacturing, so that the high aspect ratio wing model can be equipped with a control surface and a control system, which is convenient for conducting research on active aeroelastic control.

[0065] In the above, the unit cell is formed by connecting the bottom surfaces of two hexagonal pyramidal frameworks, such as Fig.10 As shown, the unit cell is composed of two hexagonal pyramid-shaped skeletons. Each of the two hexagonal pyramids has an identical bottom surface. The unit cell is obtained by splicing their identical bottom surfaces up and down. The angle between the side surfaces of the hexagonal pyramids constituting the unit cell and the central axis of the unit cell is 48°.

[0066] According to this unit cell, we can get Fig.11 The design method of the flexible wing structure with internally filled unit cells is as follows: in the geometric design software Solidworks, the wing shape is first determined according to the aerodynamic requirements of the flexible wing, and the skin that matches the shape is designed, and then the design is as follows Fig.10 The unit cells shown in the figure are arranged in an array at equal intervals to form a lattice structure. The wing skin is placed at the lattice, and the intersection of the lattice and the skin is cut to form the internal filling structure of the wing. Then the filling structure is combined with the wing skin to obtain the following: Fig.11 The preliminary structure of the flexible wing is shown.

[0067] Based on the above high aspect ratio wing, this case proposes the following optimization methods, such as Fig.12 As shown, the optimization target of this case is the natural frequency, which is affected by the mass and stiffness of the wing, and the length of the wing affects the stiffness:

[0068] Step 1, establishing a high-fidelity structural finite element model of the flexible wing model;

[0069] The flexible wing used in this embodiment is as follows Fig.13 The wing comprises a wing-body fusion part, a wing section, a control surface (3) and a winglet, the wing skin thickness is 1.5 mm, and the thickness of the internal filling cell is 1.2 mm.

[0070] A high-fidelity geometric model of the wing was created in Solidworks, and then imported into the finite element modeling software Patran. In the software, the finite element mesh of the model was accurately established according to the shape and size of the geometric model, and the mesh unit properties and material properties were assigned. The material selected was ULTEM 9085, and the anisotropic model was used to define the material.

[0071] Step 2, establishing a size optimization model of the flexible wing;

[0072] The objective function of the optimization model is the sum of the squares of the deviations between the natural frequencies of the first five modes of the flexible wing except the in-plane mode and the scaled natural frequencies of the prototype wing, which is required to be minimized, as shown in the following formula:

[0073] where i is the order, ω αi is the natural frequency of the i-th order mode obtained experimentally, that is, the natural frequency after optimization, ω mi is the natural frequency of the i-th order mode obtained by finite element, that is, the natural frequency of the prototype wing;

[0074] The design variables of the optimization model include the thickness of each unit cell filled inside the wing, the thickness of the skin, and the thickness of the winglet; among the design variables, the range of variation of the unit cell thickness is 0.6-3mm, the range of variation of the skin thickness is 0.75-3.75mm, and the range of variation of the winglet thickness is 3-15mm.

[0075] The constraints of the optimization model include stress constraints, natural frequency constraints, vibration mode constraints, and flutter speed constraints; then the upper and lower boundaries of the design variables and constraints are given;

[0076] In the constraint conditions, the upper limit of the stress constraint is set to the strength limit of the material, and the lower limit is set to 0.01MPa;

[0077] The variation range of the natural frequency constraint is 10% above and below the target natural frequency obtained by scaling the natural frequency of the physical aircraft;

[0078] The range of the vibration mode constraint is 10% above and below the target vibration mode coordinates obtained after scaling calculation of the vibration mode coordinates of the physical aircraft. The flutter velocity constraint is the damping of the flutter branch, and its boundary is that the damping is less than 0.03 at the target flutter velocity determined according to the flight conditions of the physical aircraft;

[0079] Step 3, using the optimization module of the finite element modeling software Patran, input the optimization model established in step 2 into it; then submit the finite element model to the finite element analysis software MSC.Nastran, perform optimization calculation according to the upper and lower boundaries of the given design variables and constraints in step 2, and optimize the design variables;

[0080] If the solution does not converge, reselect the upper and lower boundaries of the design variables and constraints and submit the calculation again; after the solution converges, derive the design variable values ​​obtained from the last iteration;

[0081] In this embodiment, the vibration modes of the first five modes of the optimized model except the in-plane mode are as follows: Fig.13 As shown in the figure, the modal vibration shapes of each order after optimization are consistent with those of the prototype wing;

[0082] Step 4: According to the optimized natural frequency, calculate the deviation of the natural frequency, modify the weight and length of the wing, so that the value of the corresponding dimension of the geometric model is equal to the optimized value.

[0083] The modal experiment of the large aspect ratio wing model was carried out using the hammer method. The hammer signal was added to the force window, and the time domain sample records outside the force window were weighted to zero, which can eliminate the noise that may come from the force hammer excitation channel; weighting the response signal with an exponential window can accelerate signal attenuation and reduce energy leakage. Through the time history of the sensor signal, a fast Fourier transform can be performed to obtain the frequency domain response curve of the point. The natural frequency of the system is determined with reference to the frequency domain scalar function obtained by integrating all channel data. At the natural frequency of the system, the scalar function MvMIF will dip or reach a minimum value. The scalar function and the frequency response curve at the origin under each working condition are shown as follows: Fig.14 As shown in the figure, MvMIF: modal indicator function; Frequency: frequency; Origin Point FRF: origin frequency response function; Amplitude: amplitude; Origin Data: original data; Synthetized Data: synthetic data.

[0084] The frequencies and damping ratios of each mode under different working conditions are shown in Table 1. The modal formations of each mode without winglets after mass normalization are as follows: Fig.15 shown.

[0085] Table 1 Natural frequency and damping ratio of each modal

[0086]

[0087] Accordingly, the lightweight structural dynamics design technology of the flexible wing provided in this embodiment reduces the weight of the wing section by 30.3% compared with the wing section with the same appearance but using a skeleton structure layout. By adjusting the angle between the side of the unit cell and the plane perpendicular to the bottom surface, the support material will not be generated in the unit cell during the additive manufacturing of the wing, so that the overall printing of the wing can be achieved, reducing the processing workload. The optimization model established during the design process takes into account the structural dynamics and flutter characteristics as constraints, so that the optimization results are more in line with the characteristics of the flying wing layout aircraft, and the dynamic characteristics of the aircraft can also be greatly adjusted.

[0088] There are many specific implementation ways of the present invention. The above is only the preferred implementation mode of the present invention. It should be pointed out that for ordinary technicians in this technical field, several improvements can be made without departing from the principle of the present invention. These improvements should also be regarded as the protection scope of the present invention.

Claims

1. A lightweight optimization method for a large aspect ratio wing, characterized in that: The following steps are involved: Step 1, establishing a high-fidelity structural finite element model of the flexible wing model; First, the designed high-fidelity wing geometry model is imported into the finite element software. Then, the finite element mesh of the model is accurately established according to the shape and size of the geometry model, and the mesh unit attributes and material attributes are assigned. Step 2, establishing a size optimization model of the flexible wing; The objective function of the optimization model is the first n orders of the flexible wing, where n is a positive integer. The sum of the squares of the deviations between the natural frequencies of the modes other than the in-plane modes and the scaled natural frequencies of the prototype wing is required to be minimum, as shown in the following formula: where i is the order, ω αi is the natural frequency of the i-th order mode obtained experimentally, that is, the natural frequency after optimization, ω mi is the natural frequency of the i-th order mode obtained by finite element, that is, the natural frequency of the prototype wing; The design variables of the optimization model include the thickness of each unit cell filled inside the wing, the thickness of the winglet, the thickness of the skin, and the thickness of the spar; the constraints of the optimization model include stress constraints, natural frequency constraints, vibration mode constraints, and flutter speed constraints; The upper and lower boundaries of the design variables and constraints are given. The upper and lower boundaries of the design variables are given according to the model processing technology limitations and experience. Among the upper and lower boundaries of the constraints, the stress constraint is given according to the strength of the model material; the natural frequency constraint is given according to the natural frequency of the physical aircraft corresponding to the flexible wing model after scaling calculation; the vibration mode constraint is given according to the vibration mode coordinates of the physical aircraft corresponding to the flexible wing model after scaling calculation; The flutter speed constraints are given according to the expected flight conditions; Step 3: Solve the optimization model based on finite element software and finite element model; Use the optimization module of the finite element modeling software Patran and input the optimization model established in step 2 into it; The finite element model is then submitted to the finite element analysis software MSC.Nastran, and optimization calculations are performed according to the upper and lower boundaries of the given design variables and constraints in step 2 to optimize the design variables; If the solution does not converge, reselect the upper and lower boundaries of the design variables and constraints and submit the calculation again; after the solution converges, derive the design variable values ​​obtained from the last iteration; Step 4: modify the wing geometry model according to the optimized design variable values; According to the optimized natural frequency, the deviation of the natural frequency is calculated, and the weight and length of the wing are modified so that the values ​​of the corresponding dimensions of the geometric model are equal to the optimized values.

2. The lightweight optimization method for a high aspect ratio wing according to claim 1, characterized in that: The interior of the wing is filled with a lattice structure composed of unit cells, the lattice structure is covered with a skin, and the unit cell is formed by connecting the bottom surfaces of two hexagonal pyramid-shaped frames.

3. The lightweight optimization method for a high aspect ratio wing according to claim 1, characterized in that: The wing comprises a wing-body fusion part (1), a wing section (2) and a winglet (4) which are connected in sequence, one side of the wing-body fusion part (1) is used to connect to the fuselage of the aircraft, and the other side is fixedly connected to the winglet (4) through two wing sections (2), each wing section (2) is respectively equipped with two control surfaces (3), and a steering gear for controlling the rotation of the control surface (3) is installed inside the wing section (2); The wing section (2) and the control surface (3) are processed by a 3D printer, the interior of which is filled with unit cells to form a lattice structure, and the lattice structure is covered with a skin; The wing-body fusion portion (1) and the wing section (2), adjacent wing sections (2), and the wing section (2) and the winglet (4) are all connected via a separation surface connection structure.

4. The lightweight optimization method for a high aspect ratio wing according to claim 3, characterized in that: Two adjacent structural members are connected by means of the separation surface connection structure, wherein the separation surface connection structure comprises: a slot (12) along the chord direction of the wing is provided on one of the structural members, and a ridge (13) along the chord direction of the wing is provided on the other structural member, and after the ridge is assembled in the slot along the chord direction, it is locked by means of a fixing screw vertically arranged through the ridge (13).

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