A lightweight optimization method for high aspect ratio airfoils

Abstract

This invention discloses a lightweight optimization method for high aspect ratio airfoils, involving fields such as structural dynamics and aircraft structural design. It comprehensively considers structural dynamics and flutter characteristics, solving the problems of limited adjustment space for dynamic characteristics and a limited number of optimization methods in the lightweight design of flexible airfoils. The steps are as follows: Step 1, establish a high-fidelity structural finite element model of the flexible airfoil; Step 2, establish a dimensional optimization model of the flexible airfoil; Step 3, solve the optimization model based on finite element software and the finite element model; Step 4, modify the airfoil geometry model according to the optimized design variable values. The optimization model established in this invention considers structural dynamics and flutter characteristics as constraints, making the optimization results more consistent with the characteristics of flying wing aircraft and facilitating further improvement of the aircraft's structural dynamic characteristics.

Description

Technical Field

[0001] This invention relates to the fields of structural dynamics and aircraft structural design, and specifically to the structural dynamics design technology of lightweight flexible wings. Background Technology

[0002] In recent years, with the aviation industry's increasing demands for advanced aircraft in terms of aerodynamics and stealth performance, flying wing aircraft have received widespread attention. These aircraft typically employ flexible wing structures with a high aspect ratio, exhibiting excellent lift-drag characteristics. Simultaneously, the use of lightweight materials in their manufacture effectively reduces structural weight, thus decreasing fuel consumption. However, due to the light weight and high flexibility of these wings, under flight loads, they undergo significant bending and torsional deformations, easily leading to complex structural dynamics problems, such as flutter. Therefore, the impact of structural dynamics characteristics must be considered when designing flexible wing structures, presenting new challenges to the lightweight design of flexible wings.

[0003] Currently, when designing lightweight internal structures for flexible airfoil models, academia and industry typically employ a skeletal structure layout that includes components such as spars, longitudinal walls, stiffeners, and ribs. The shapes and locations of these components are generally fixed, thus limiting the design scope for lightweighting and hindering adjustments to the aircraft's dynamic characteristics.

[0004] Researchers and engineers often employ optimization methods when designing lightweight flexible wings. A common optimization approach in existing research involves first performing aerodynamic optimization on the wing, then optimizing its dimensions while considering aerodynamic loads and structural strength constraints, with weight reduction as the primary goal. This yields an optimized result with good aerodynamic and strength characteristics. However, this method pays less attention to structural dynamics and easily overlooks the impact of flutter on flexible wings. Furthermore, the finite element models of wing structures used in existing optimization efforts are generally simplified to some extent, potentially leading to discrepancies between the optimization results and the actual models.

[0005] Furthermore, with the continuous development of materials and manufacturing technologies, additive manufacturing technology, with its advantages of simple processing steps and fast prototyping, has become one of the important means of future aircraft design and manufacturing. For flying wing aircraft, additive manufacturing can precisely control the shape and position of the internal stress structure, providing convenience for lightweight structural design and adjustment of the mechanical properties of the model.

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

[0007] To address the above problems, this invention proposes a lightweight optimization method for high aspect ratio wings, which comprehensively considers structural dynamics and flutter characteristics, and solves the problems of limited space for dynamic characteristic adjustment and few optimization methods in the lightweight design of flexible wings.

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

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

[0010] First, import the designed high-fidelity wing geometric model into the finite element software. Then, accurately create the finite element mesh of the model according to the shape and size of the geometric model, and assign mesh element properties and material properties.

[0011] Step 2: Establish a size optimization model for the flexible wing;

[0012] The objective function of the optimization model is the sum of squares of the deviations between the natural frequencies of the first n modes of the flexible wing (excluding in-plane modes) and the natural frequencies of the scaled-down prototype wing, and is required to be minimized, as shown in the following formula:

[0013] Where i is the order, ω αi Let ω be the natural frequency of the i-th mode obtained from the experiment, i.e., the optimized natural frequency. mi Let be the natural frequency of the i-th mode obtained by finite element method, i.e., the natural frequency of the prototype wing;

[0014] The design variables of the optimization model include the thickness of each unit cell filling the wing interior, the thickness of the winglets, the thickness of the skin, and the thickness of the spars; the constraints of the optimization model include stress constraints, natural frequency constraints, mode shape constraints, and flutter velocity constraints.

[0015] For design variables and constraints, upper and lower boundaries must be given. The upper and lower boundaries of design variables are given based on model manufacturing process limitations and experience. Among the upper and lower boundaries of constraints, stress constraints are given based on the strength of the model material; natural frequency constraints are given after scaling calculation based on the natural frequencies of the actual aircraft corresponding to the flexible wing model; mode shape constraints are given after scaling calculation based on the mode shape coordinates of the actual aircraft corresponding to the flexible wing model; flutter velocity constraints are given based on the expected flight conditions.

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

[0017] 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, and perform optimization calculations according to the upper and lower boundaries of the design variables and constraints given in step 2 to 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, export the design variable values ​​obtained from the last iteration.

[0019] Step 4: Modify the wing geometry model based on the optimized design variable values;

[0020] Based on the optimized natural frequencies, the deviation of the natural frequencies is calculated, and the weight and length of the wing are modified so that 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, and the lattice structure is covered with a skin. The unit cells are formed by two hexagonal frustum-shaped skeletons with their bottom surfaces connected.

[0022] The wing includes a wing-body blend 1, a wing segment 2, and a winglet 4 connected in sequence. One side of the wing-body blend 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 segments 2. Each wing segment 2 is equipped with two control surfaces 3, and a servo motor for controlling the rotation of the control surfaces 3 is installed inside the wing segment 2.

[0023] The wing segment 2 and the control surface 3 are manufactured by a 3D printer. Their interiors are filled with unit cells to form a lattice structure, and a skin is wrapped around the lattice structure.

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

[0025] The two adjacent structural components are connected by the separation surface connection structure. The separation surface connection structure is as follows: a slot 12 along the chord direction of the wing is provided on one structural component, and a protrusion 13 along the chord direction of the wing is provided on the other structural component. After the protrusion 13 is assembled into the slot along the chord direction, it is locked by a vertically arranged fixing screw that passes through the protrusion 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 and thin-walled, which can effectively reduce the weight of the wing. Furthermore, if the thickness of each unit cell is designed individually, the mass and stiffness distribution in various parts of the wing can be changed, thus giving the aircraft's dynamic characteristics a large adjustment range.

[0028] Second, when the model is processed using additive manufacturing, by adjusting the angle between the side of the cell unit and the plane perpendicular to the bottom surface, support material will not be generated in the cell unit when printing along the chord of the wing, thus facilitating the overall printing of the wing and reducing the amount of processing work.

[0029] Third, using a high-fidelity finite element model for analysis, and selecting an anisotropic model that better reflects the actual properties of additive manufacturing materials to define material properties, can reduce the error between simulation results and physical model representation.

[0030] Fourth, 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. Attached Figure Description

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

[0032] Figure 2 This is a top view of the structural layout of a high aspect ratio airfoil model;

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

[0034] Figure 4 These are the top and side views of the control surface.

[0035] Figure 5 This is an assembly diagram of the connection structure between the control surfaces and the wing section;

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

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

[0039] Figure 9 This is a schematic diagram of a round shaft with a spring used in the mechanism connecting the control surfaces and the wing section;

[0040] Figure 10 This is a schematic diagram of the external shape of the unit cell structure provided by the present invention;

[0041] Figure 11 It is based on Figure 10 A schematic diagram of a flexible airfoil structure obtained from a medium-cell unit cell.

[0042] Figure 12 This is a flowchart of a dynamic design method for a lightweight flexible airfoil structure.

[0043] Figure 13 These are the mode shapes of the first 5 modes of the optimized model, excluding the in-plane modes.

[0044] Figure 14 It shows the scalar function and frequency response curve at the origin for each operating condition;

[0045] Figure 15 These are the modal array diagrams of each order without winglets after mass normalization;

[0046] In the diagram, 1 is the wing-body fusion section, 2 is the wing section, 3 is the control surface, 4 is the winglet, 5 is the servo canopy, 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 reinforcing rib, 12 is the groove, 13 is the protrusion, 17 is the aluminum alloy round shaft, and 18 is the spring. Detailed Implementation

[0047] To clearly illustrate the technical features of this patent, the following detailed description is provided through specific embodiments and in conjunction with the accompanying drawings.

[0048] To ensure that additively manufactured high-aspect-ratio wings can be easily connected and assembled, and to accurately and efficiently complete the installation of various parts of the high-aspect-ratio wing, the wing, as... Figure 1 and Figure 2 As shown, it includes a wing-body blending section 1, two wing sections 2, four control surfaces 3, and a winglet 4. The overall model has a span of 1.6m.

[0049] One side of the wing-body fusion section 1 is used to connect to the fuselage of the aircraft, and the other side is fixedly connected to the winglets 4 via two wing sections 2. Each wing section 2 is equipped with two control surfaces 3, and a servo motor for controlling the rotation of the control surfaces 3 is installed inside the wing section 2.

[0050] Among them, the internal use of the wing is as follows Figure 3 The wing's internal filling structure is obtained by filling a lattice structure composed of unit cells, placing the wing skin at the lattice, and cutting off the portion where the lattice intersects with the skin. The wing segment 2 and the control surface 3 are manufactured using additive manufacturing processes, and after forming, they contain unit cells and skin.

[0051] Each wing segment 2 has two servo nacelles in the middle of its lower surface, such as Figure 2 As shown, the interior is used to install servo motors that control the deflection of the control surfaces, and the exterior is fixedly installed with a servo motor cover 5 for sealing.

[0052] like Figure 5 As shown, the trailing edge of the wing section 2 has a receiving groove for accommodating the control surface 3. The inner wall of the receiving groove has a spanwise mounting hole 9 with a diameter of 10 mm. Each hole has a rolling bearing 10 for rotatably connecting with the shaft head of the control surface 3.

[0053] The shape of 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, and taking the part including the trailing edge. The front part of the control surface 3 has a spanwise hinge hole 6 with a diameter of 6.8 mm; after the pivot is inserted into the hinge hole 6 and the mounting hole 9, the control surface 3 is rotatably connected to the receiving groove of the airfoil section 2.

[0054] There is a tapered limiting hole 7 at each end of the control surface 3. The limiting hole 7 points to the rotating shaft. After the limiting pin is installed in the limiting hole 7, the control surface 3 and the rotating shaft are fixed relative to each other. There are also three threaded holes 8 in the middle of the control surface 3, which are arranged in a triangle for installing the transmission mechanism of the servo motor to accept the control of the servo motor to flip up and down.

[0055] The winglet 4 is divided into two semi-winglets vertically from the chord, as shown in the image. Figure 6a , 6b As shown, bolt holes are arranged around the side edges of the two winglets for connecting them. The two winglets have internal reinforcing ribs 11, as shown... Figure 6a As shown, one type of semi-wing has a transition section on its outer wall for connecting wing segments.

[0056] Furthermore, this embodiment also includes a connection structure between the control surface and the wing section and its assembly schematic diagram, as shown below. Figure 5 As shown. The rotating shaft comprises two identical aluminum alloy round shafts 17 and a spring 18, as... Figure 9 As shown. The diameter of the round shaft 17 is 6mm. The aluminum alloy round shaft 17 has a long groove for connecting the limiting pin. The spring 18 is fixedly connected between the two aluminum alloy round shafts 17.

[0057] When installing the control surface 3, first insert the two aluminum alloy round shafts 17, each equipped with a spring 18, into the hinge holes 6 along the span of the control surface 3, and press both ends of the aluminum alloy round shafts 17 by hand so that they are entirely inside the through holes; then move the control surface 3 so that the axis of the hinge holes 6 is aligned with the axis of the control surface 3. Figure 5 Align the axes of the mounting holes 9 in the middle, then release the handle so that the two aluminum alloy round shafts 17 extend into the mounting holes 9 along the spanwise direction at both ends of the wing section mounting part under the restoring force of the spring 18; finally, insert the limiting pins into the limiting holes 7 at both ends of the control surface 3 surface, and insert the pins into the long grooves on the surface of the aluminum alloy round shaft 17 at the same time to complete the installation of the control surface.

[0058] In this case, adjacent wing segments 2 are connected by a separation surface connection structure, and the connection structure and its assembly effect are shown in Figures 7 and 8.

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

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

[0061] In this way, after the chordally oriented protrusion 13 and slot 12 are assembled, they are locked in place by the vertical fixing screws. Finally, the separation surface connection structure can achieve full positioning of the structural components on both sides of the separation surface in the chord, span, and vertical directions. This not only simplifies assembly but also ensures precise positioning and allows for repeated disassembly.

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

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

[0064] The aforementioned separation surface connection structure reduces the number of fasteners used in wing segment connections, simplifying wing segment installation; furthermore, wing segments connected using this method can be repeatedly disassembled and reassembled. The control surface and wing segment connection structure provided by this invention is relatively simple to assemble, requires fewer separation surfaces, and can be achieved through additive manufacturing. This allows high-aspect-ratio wing models to incorporate control surfaces and control systems, facilitating research into aeroelastic active control.

[0065] In the above, the unit cell is formed by two hexagonal frustum-shaped frameworks with their base surfaces joined together, such as... Figure 10 As shown, the unit cell is composed of two hexagonal frustum-shaped frameworks. Each of the two hexagonal frustums has an identical base. The unit cell is obtained by piecing together their identical bases one above the other. The angle between the side of the hexagonal frustum that makes up the unit cell and the central axis of the unit cell is 48°.

[0066] Based on this unit cell, we can obtain, as follows: Figure 11 The flexible wing structure with internally filled unit cells shown is designed as follows: In the geometric design software Solidworks, the wing shape is first determined according to the aerodynamic requirements of the flexible wing, and a skin conforming to the shape is designed. Then, a structure like... Figure 10 The unit cells shown are arranged equidistantly in an array to form a lattice structure. The wing skin is placed at the lattice points, and the portion where the lattice intersects with the skin is cut off to form the internal filling structure of the wing. Then, the filling structure is combined with the wing skin to obtain the structure shown. Figure 11 The preliminary structure of the flexible wing shown.

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

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

[0069] The flexible wing used in this embodiment is, for example... Figure 13 As shown. The wing includes a wing-body fusion section, wing section, control surfaces (3) and winglets. The skin thickness of the wing is 1.5 mm, and the thickness of the internal filling cells 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 created according to the shape and size of the geometric model, and the mesh element 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: Establish a size optimization model for 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 (excluding the in-plane modes) and the natural frequencies of the scaled-down prototype wing, and the objective function is to minimize this sum, as shown in the following formula:

[0073] Where i is the order, ω αi Let ω be the natural frequency of the i-th mode obtained from the experiment, i.e., the optimized natural frequency. mi Let be the natural frequency of the i-th mode obtained by finite element method, i.e., the natural frequency of the prototype wing;

[0074] The design variables of the optimization model include the thickness of each cell unit filling the wing interior, the skin thickness, and the winglet thickness; among the design variables, the cell unit thickness varies from 0.6 to 3 mm, the skin thickness varies from 0.75 to 3.75 mm, and the winglet thickness varies from 3 to 15 mm.

[0075] The constraints of the optimization model include stress constraints, natural frequency constraints, mode shape constraints, and flutter velocity 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 material's strength limit, and the lower limit is set to 0.01 MPa;

[0077] The range of variation 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 variation range of the mode shape constraint is 10% above and below the target mode shape coordinates obtained after scaling the mode shape coordinates of the actual 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 actual aircraft.

[0079] Step 3: Use the optimization module of the finite element modeling software Patran to input the optimization model established in Step 2; then submit the finite element model to the finite element analysis software MSC.Nastran, and perform optimization calculations according to the upper and lower boundaries of the design variables and constraints given in Step 2 to 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, export the design variable values ​​obtained from the last iteration.

[0081] In this embodiment, the mode shapes of the first 5 modes of the optimized model, excluding the in-plane modes, are as follows: Figure 13 As shown, the optimized mode shapes are consistent with those of the prototype wing;

[0082] Step 4: Based on the optimized natural frequency, calculate the deviation of the natural frequency, and modify the weight and length of the wing so that the corresponding dimensions of the geometric model are equal to the optimized values.

[0083] Modal experiments on a high aspect ratio airfoil model were conducted using the hammer impact method. A force window was applied to the hammer impact signal, and time-domain samples outside the force window were weighted and set to zero to eliminate noise that might originate from the hammer excitation channel. Using an exponential window to weight the response signal accelerated signal attenuation and reduced energy leakage. The frequency response curve at a given point was obtained by performing a Fast Fourier Transform (FFT) on the time history of the sensor signals. The natural frequency of the system was determined using a frequency-domain scalar function obtained by integrating data from all channels. At the system's natural frequency, the scalar function MvMIF will exhibit a dip or a minimum value. The scalar function and frequency response curves at the origin for each operating condition are shown below. Figure 14 As shown in the figure, MvMIF: Modal Indication Function; Frequency: Frequency; Origin Point FRF: Origin Frequency Response Function; Amplitude: Amplitude; Origin Data: Raw Data; Synthetized Data: Synthetic Data.

[0084] The frequencies and damping values ​​corresponding to each mode under different operating conditions are shown in Table 1 below. The mode patterns of each mode excluding winglets after mass normalization are as follows: Figure 15 As shown.

[0085] Table 1 Natural frequencies and damping ratios of each modal mode.

[0086]

[0087] Accordingly, the flexible wing lightweight structural dynamics design technology provided in this embodiment reduces the wing section weight by 30.3% compared to a wing section with the same shape but using a skeleton structure layout. By adjusting the angle between the side of the cell unit and the plane perpendicular to the bottom surface, no supporting material is generated within the cell unit during wing additive manufacturing, thus enabling the integral printing of the wing and reducing the amount of processing work. The optimization model established during the design process considers structural dynamics and flutter characteristics as constraints, making the optimization results more consistent with the characteristics of flying wing aircraft and allowing for significant adjustments to the aircraft's dynamic characteristics.

[0088] There are many specific ways to implement this invention. The above description is only a preferred embodiment of this invention. It should be noted that for those skilled in the art, several improvements can be made without departing from the principle of this invention, and these improvements should also be considered within the scope of protection of this invention.

Claims

1. A lightweight optimization method for a high aspect ratio airfoil, characterized in that, The following steps are involved: Step 1: Establish a high-fidelity structural finite element model of the flexible wing model; First, import the designed high-fidelity wing geometric model into the finite element software. Then, accurately create the finite element mesh of the model according to the shape and size of the geometric model, and assign mesh element properties and material properties. Step 2: Establish a size optimization model for the flexible wing; The objective function of the optimization model is the sum of squares of the deviations between the natural frequencies of the first n modes of the flexible wing (excluding in-plane modes) and the natural frequencies of the scaled-down prototype wing, and is required to be minimized, as shown in the following formula: Where i is the order, ω αi Let ω be the natural frequency of the i-th mode obtained from the experiment, i.e., the optimized natural frequency. mi Let be the natural frequency of the i-th mode obtained by finite element method, i.e., the natural frequency of the prototype wing; The design variables of the optimization model include the thickness of each unit cell filling the wing interior, the thickness of the winglets, the thickness of the skin, and the thickness of the spars; the constraints of the optimization model include stress constraints, natural frequency constraints, mode shape constraints, and flutter velocity constraints. The upper and lower boundaries of the design variables and constraints must be given. The upper and lower boundaries of the design variables are given according to the model manufacturing process limitations and experience. Among the upper and lower boundaries of the constraints, the stress constraints are given according to the strength of the model material; the natural frequency constraints are given after scaling calculation based on the natural frequencies of the actual aircraft corresponding to the flexible wing model; and the mode shape constraints are given after scaling calculation based on the mode shape coordinates of the actual aircraft corresponding to the flexible wing model. Flutter velocity constraints are given based on 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 to input the optimization model established in step 2; 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 design variables and constraints given 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, export the design variable values ​​obtained from the last iteration. Step 4: Modify the wing geometry model based on the optimized design variable values; Based on the optimized natural frequencies, the deviation of the natural frequencies is calculated, and the weight and length of the wing are modified so that the corresponding dimensions of the geometric model are equal to the optimized values.

2. The lightweight optimization method for a high aspect ratio airfoil according to claim 1, characterized in that, The interior of the wing is filled with a lattice structure composed of unit cells, and the lattice structure is covered with a skin. The unit cells are formed by two hexagonal frustum-shaped skeletons with their bottom surfaces connected.

3. The lightweight optimization method for a high aspect ratio airfoil according to claim 1, characterized in that, The wing includes a wing-body fusion section (1), a wing segment (2), and a winglet (4) connected in sequence. One side of the wing-body fusion section (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 segments (2). Each wing segment (2) is equipped with two control surfaces (3), and a servo motor for controlling the rotation of the control surfaces (3) is installed inside the wing segment (2). The wing segment (2) and the control surface (3) are manufactured by a 3D printer. Their interiors are filled with unit cells to form a lattice structure, and a skin is wrapped around the lattice structure. The wing-body fusion section (1) and the wing segment (2), adjacent wing segments (2), and the wing segment (2) and the winglet (4) are all connected by a separation surface connection structure.

4. The lightweight optimization method for a high aspect ratio airfoil according to claim 3, characterized in that, The two adjacent structural components are connected by the separation surface connection structure. The separation surface connection structure is as follows: a slot (12) along the chord direction of the wing is provided on one structural component, and a protrusion (13) along the chord direction of the wing is provided on the other structural component. After the protrusion is assembled into the slot along the chord direction, it is locked by a vertically arranged fixing screw that passes through the protrusion (13).

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