In the flight process, the morphing aircraft can adapt to various mission requirements in real time by partially or wholly changing the shape of the aircraft, and can maintain the optimal performance and efficiency in various environments, effectively improve the flight performance of the aircraft, expand the flight envelope and improve the flight efficiency^[1].

As the most important part of aircraft aerodynamic force, the aerodynamic performance of the wing directly affects the performance and flight efficiency of the aircraft, so designing a wing whose shape can be adaptively adjusted according to the flight state and environment will greatly improve the performance and flying quality of the aircraft. Different from the non-smooth continuous airfoil camber change realized by high-lift mechanisms such as leading edge slats, trailing edge flaps and ailerons in modern aircraft, the development of morphing wings with continuous chordwise camber change is one of the hot research directions of morphing aircraft at present^[2]. For transport aircraft, on the one hand, the chordwise variable camber wing can maintain the optimal lift-drag ratio in real time according to the real-time state of the aircraft, save fuel and increase the range; On the other hand, it can also be combined with the design of control law to achieve the benefits of aircraft load alleviation, aeroelastic tailoring, noise reduction and structural weight reduction. At present, the research of chordwise variable camber wing mainly focuses on the technology of variable camber at the leading edge and the trailing edge.

At present, all countries in the world are competing to take the chordwise variable camber wing technology as one of the key technical fields to be studied. In 1981, NASA's Dryden Institute developed the "Mission Adaptive Wing" and conducted modification tests on the F-111A fighter bomber^[3]. The leading and trailing edges of the modified wing are each composed of a portion of active surfaces made of glass fiber reinforced plastic, which can be manipulated to change the camber and twist angle of the wing according to different flight conditions. Later, it was not really applied because of its heavy and complex mechanical structure. Since 2010, NASA and Boeing have cooperated in the project of "Continuous Variable Camber Trailing Edge Flap System", which is dedicated to the development of a new three-section smooth variable camber wing trailing edge driven by memory alloy and distributed motors.The flap system can achieve the optimal lift-drag ratio and save fuel consumption in multi-mission conditions, but the system is still in the prototype development stage^[4]. Lu et al., Miller et al., and Kota et al. Completed the adaptive trailing-edge flap design using a compliant mechanism, and tested the camber change ability of the adaptive trailing-edge flap during flight through dynamic flight tests on a test aircraft^[5][6][7]. In recent years, NASA has designed and manufactured a cellular array combined variable camber wing with ultra-light structure and appropriate stiffness and strength by using the idea of cellular array structure^[8]. Sinapius et al. Proposed the concept of "finger" type deformation, based on the traditional mechanical structure scheme, through the internal four-link design, to achieve single-degree-of-freedom drive^[9]. Woods et al. Proposed a fishbone scheme to realize variable camber design based on the principle of bionics^[10]. Fisher designed and manufactured a structure with zero Poisson's ratio, and manufactured the fishbone skeleton and the honeycomb structure inside the skin through 3D printing technology, which solved the problem that the flexible skin and the skeleton could not be deformed and matched^[11]. The European "Clean Sky" project has developed three morphing structures for regional airliners: drooped nose, multi-function flap and adaptive winglet, aiming to adjust the shape of the aircraft in real time according to flight conditions to achieve optimal aerodynamic efficiency, and evaluated the effectiveness of the three morphing devices on laminar flow wings^[12]. Snow et al. Manufactured a flexible wing. The whole wing is an integral structure, which is not detachable. It is made by 3D printing technology. The performance and stability of the wing are discussed, and the prospect is put forward^[13]. Sofla et al. Studied the method of wing bending actuation using two non-actuated deformation positions maintained by one-way shape memory alloy^[14]. Yang et al. Used shape memory alloy as an actuator to drive the variable camber wing of a small UAV, and analyzed the aerodynamic performance and aeroelasticity of the wing under this driving mode^[15]. Vos et al. Described how to use piezoelectric actuators to manipulate the camber distribution in a deformable wing structure to achieve roll control on a small UAV^[16]. Mkhoyan et al. Designed and developed an autonomous morphing wing and proposed a new distributed morphing concept, which can control the spanwise lift distribution of the wing to ensure the optimal performance of the wing^[17]. Jenett et al. Designed a morphing wing based on discrete modular units, which is composed of discrete units and can perform continuous lateral torsional deformation to improve the rolling efficiency of the wing^[18]. Zhang et al. Established an aeroelastic model of a cambered morphing wing, and used its chordwise size and flexibility to study the flutter critical speed^[19]. By comparing the quasi-steady aerodynamic model with the unsteady aerodynamic model, it is determined that the quasi-steady aerodynamic model is more conservative in predicting the flutter speed.

In domestic research, Yang Chao and others have conducted preliminary research on active aeroelastic wing and aeroelastic response load mitigation^[20]. The morphing wing driven by ultrasonic motors developed by Nanjing University of Aeronautics and Astronautics uses distributed ultrasonic motors to drive a rigid backyard rotatable lever to realize the chordwise deformation of the wing^[21]. Zhang Yinxuan et al. Designed a flexible skin structure, which is composed of flexible honeycomb and elastic adhesive film, and has good deformation capacity in plane and certain bearing capacity out of plane^[22]. Yang Zhichun et al. Connected the multi-piece ribs through the connecting rod slider and the sliding hinge group, designed the trailing edge structure of the wing with variable camber, and systematically studied its deflection configuration, controlled kinematics and aerodynamic characteristics^[23]. Chen et al. And Cold Jinsong et al. Developed a chordwise variable camber wing structure based on shape memory alloy and pneumatic tendons^[24][25]. Chen Xiu et al. Proposed a fully flexible wing structure design scheme based on the continuous topology optimization method, combined the topology optimization method with the wing structure design, and studied the optimization design of the leading and trailing edge variable camber structure of the wing respectively^[26]. Xu Junheng et al proposed a variable camber wing based on cross reed hinge, established its mechanical model and optimized its design. The test results show that the wing has large deformation capacity and can realize continuous chordwise variable camber^[27]. Wang Yu et al. Designed a structure with a curved beam and a planar disk driving the trailing edge deflection by using a honeycomb structure flexible material with zero Poisson's ratio as a flexible skin, and studied its multidisciplinary design and optimization method^[28]. Wei Rukai et al. carried out a numerical study on gust response alleviation based on the variable camber trailing edge, and found that the variable camber trailing edge has greater potential in gust alleviation control than the traditional hinged control surface^[29].

Compared with the related work of variable camber wing in recent years, new driving technologies such as shape memory alloy, piezoelectric ceramics and ultrasonic motor have become the focus of attention of scholars at home and abroad. Among them, shape memory alloy has the advantages of large strain and large stress, but the heating and cooling time of metal wire is long, the driving ability is greatly affected by environmental factors, and the driving performance is unstable and inefficient. At the same time, the fatigue life of the metal wire will seriously restrict the service time of the actuator. Although piezoelectric materials have the advantages of high sensitivity and strong designability, their driving speed is slow and the output deformation is limited, so they are more suitable for small morphing aircraft. Ultrasonic motor has the advantages of small size and strong controllability, but its output driving force is often small because of its small size, and it needs multi-motor drive when it is used, which increases the weight of the structure and reduces the aerodynamic benefits brought by the compliant structure.

Compliant mechanism is a kind of mechanism in which the input load realizes the displacement change of the output end through its own elastic deformation and stores part of the input energy in the form of strain energy. Through the compliant mechanism, the structure can naturally realize continuous and smooth deformation in the target direction, reduce the number of parts required by the structure, replace the kinematic pair of the traditional mechanical transmission mode, reduce the wear of the mechanism, and improve the driving efficiency.

The main difficulties in the design of chordwise variable camber wing are structural design, skin matching design and actuation design. At the beginning of the structural design, the chordwise variable camber configuration of the wing is particularly important, which determines whether the deformed wing meets the design requirements, whether the aerodynamic efficiency is optimal, and whether the material stress level is reasonable under different deflection conditions.

In view of this, the compliant mechanism is introduced into the design of the chordwise variable camber wing, and a prototype of the trailing edge of the wing with rigid-flexible hybrid deflection is designed, considering that the elastic allowable deformation capacity of the material can meet the requirements of both the high-lift index and the optimal rent-increasing characteristics. Firstly, the proportion of the rigid part and the flexible part in the rigid-flexible trailing edge deflection section is determined by comprehensively considering the driving mode, the structural material, the deflection target angle and other factors; Secondly, the bending configuration of the airfoil is described by the geometric parameterization of the mean camber line of the airfoil. Thirdly, taking the lift-drag ratio of the airfoil as the optimization objective, the optimal camber configuration is determined by genetic algorithm combined with XFOIL software, and the precise lift-drag characteristics of the optimized configuration are confirmed by computational fluid dynamics (CFD), and the lift coefficient, lift-drag ratio and flow field characteristics of the traditional rigid deflection trailing edge and the rigid-flexible hybrid trailing edge under different flight conditions are compared. Then, the structural optimization principle and process of the compliant rib and the design method of the sliding skin are introduced. Finally, the deflection section of the rigid-flexible hybrid trailing edge wing was manufactured and assembled, and the preliminary deformation capability test was completed.

The flap installed on the trailing edge of the wing usually accounts for 25% to 35% of the total chord of the wing section. Taking the flap system of a medium-range medium transport aircraft as the design reference, the total chord length of the wing is determined to be 3. 33 m, the variable camber section of the trailing edge is 1 m, accounting for 30%, and the target deflection angle of the variable camber trailing edge is 30 °. The NACA0012 airfoil is selected as the basic reference airfoil for design and optimization.

In the design of the trailing edge of a variable camber wing, the input point, the structural form, and the output point constitute the basic design elements. The input point is the output position of the driver, which can be divided into two types of driving methods: centralized single-point driving and distributed driving. The structural form is obtained by topology, size, shape and other optimization methods. The output point set of the trailing edge of the chordwise variable camber is the outer contour line of the rib structure. This paper focuses on the aerodynamic characteristics optimization of the rigid-flexible hybrid variable camber trailing edge, the rigid-flexible combination design of deformable ribs and the feasibility verification of the principle. The verification of ignoring chord length change, skin matching, and load bearing capacity is not involved.

In this paper, aviation 7075-T6 duralumin alloy is selected as the design material of the variable body rib. The finite element analysis shows that if the whole deformation section of the trailing edge is designed as a compliant structure, the stress of the upper and lower flanges will exceed the allowable stress of the material when the structure is bent down to the target angle. Further analysis of the continuous variable camber requirement of the airfoil trailing edge camber line shows that the wing trailing edge deformation is similar to the cantilever beam bending deformation under load, and the initial section of the cantilever beam trailing edge camber line deformation and the wing tip section show similar equal slope distribution^[19]. Therefore, considering the elastic allowable deformation capacity of the rib material, taking into account the requirements of the high-lift index and the realization of the optimal rent-raising characteristics, this paper proposes a rigid-flexible hybrid deformable rib design scheme as shown in Figure 1. The variable camber section at the trailing edge of the wing is arranged alternately according to the rigid, flexible and follow-up sections, which account for 30%, 40% and 30% of the whole trailing edge of the wing, respectively.

For the trailing edge of a flexible wing, if the target deflection angle of the variable camber section is fixed, such as 30 °, there are many options for the deflection curve configuration to actually achieve this deflection target, and the two points of the mean camber line can be formed by different curves. Therefore, it is necessary to determine the optimal airfoil variable camber configuration based on the aerodynamic performance of the wing, so that the rigid-flexible connection has good continuity, and the overall has the optimal lift-drag ratio characteristics.

The trailing edge of the variable camber configuration is divided into rigid deflection section, flexible section and follow-up section, and its geometric characteristics are described by the mean camber line, which simplifies the relatively complex airfoil parameterization problem into a curve parameterization problem. In order to ensure the continuity and smoothness of the camber change and increase the diversity of the camber change description, the cubic spline equation is used to describe the mean camber line of the airfoil:

Where :a₁~a₄ are the coefficients of the cubic spline equation; (x_s,y_s) is the coordinate of the initial deflection point of the mean arc; (x_e,y_e) is the coordinate of the trailing edge point of the mean camber line; L is the length of the mean camber line of the variable camber trailing edge section; θ₂ is the equivalent deflection angle of the variable camber trailing-edge segment.

The camber line of the airfoil in the rigid-flexible hybrid structure is shown in Figure 2.

In Fig. 2, the six points m₀~m₅ are all located on the camber line of the airfoil, where m₀ is the initial point of the variable camber section of the wing, let it be the coordinate origin, and θ₁ is the rigid deflection angle. m₅ is the trailing edge point, m₁ and m₄ are the joint points of the rigid-flexible joint respectively. In order to ensure the smooth continuity of the rigid segment and the flexible segment, the tangent direction of the flexible segment equation at the coordinate of m₁ point should be consistent with the slope of the rigid rotating segment.The m₂ and the m₃ of the flexible section are the points with the abscissa of 400 mm and 600 mm respectively, which are located on the middle arc line of the flexible section. Substitute the coordinates of each point into the formula (1) to obtain

Where: (UNK 1, y_{m_i}) are the coordinates of the m_i point, I = 0,1, …, 5; a, B, C, d are the coefficients; l_{m_1{m}_4} is the arc length between the m₁ point and the m₄ point; l_{m_4{m}_5} is the arc length between the point m₄ and the point m₅.

Through the simultaneous simplification of the equations in Equation (2), the parameters of the mean camber curve can be transformed into equations about the rigid deflection angle θ₁ and the y coordinate of the point m₂. Therefore, the absolute values of the rigid deflection angle θ₁ and the y coordinate of the point m₂ are taken as the design variables of the optimization problem. Because the advantage of the chordwise variable camber airfoil lies in the improvement of the lift-drag characteristics compared with the traditional airfoil, the evaluation function S of the optimal variable camber configuration is the maximum lift-drag ratio of the airfoil:

Where :C_L is the lift coefficient; C_D is the drag coefficient.

The lift-drag ratio of the corresponding variable camber configuration is calculated by using the genetic algorithm. The wing trailing edge target deflection angle of 30 ° is the design maximum deflection angle, which corresponds to the landing (maximum deflection angle) state of the aircraft, so the flow field parameters are set as the aircraft landing condition, as shown in Table 1.

The calculation process is shown in Figure 3.

The optimization process sets the following constraints: considering the material strength limitation, the maximum allowable equivalent chordwise camber deflection angle produced by the compliant segment and the follower segment is 15 °. Corresponding to the constraint condition, the final optimal result is θ₁=20.06°, y_{m_2} = − 163.87 according to the evaluation function calculation in formula (3). After normalizing the coordinates of the mean camber line, the airfoil contour line corresponding to the mean camber line is drawn, and the final variable camber configuration is obtained as shown in Fig. 4.

The optimized rigid-flexible hybrid deformable rib camber line and airfoil contour line are compared with the cantilever trailing edge deformation camber line and rib, as shown in Figure 5. The maximum profile difference of the variable camber trailing edge airfoil is only 1.1% of the airfoil chord, which is 3.5% of the characteristic chord of the variable camber trailing edge.

Based on the optimized rigid-flexible hybrid variable trailing edge airfoil and the traditional rigid deflection airfoil, the flow field characteristics and aerodynamic characteristics are compared and analyzed by computational fluid dynamics analysis.

Taking the rigid-flexible mixed variable camber trailing edge airfoil with 30 ° deflection as an example, the grid model is shown in Fig. 6.

The k-ω SST turbulence model is used in the calculation, and the fluid medium is ideal gas with a temperature of 300 K, a reference standard atmospheric pressure of 101325 Pa, a Mach number of 0.2, and a reference chord length of 1000 mm.

Fig. 7 shows the variation of lift coefficient with angle of attack for two different deflection modes when the trailing edge deflection angle is 30 °. It can be seen from Fig. 7 that under the same trailing edge deflection angle, the lift coefficient of the rigid-flexible hybrid deflection mode is always larger than that of the rigid-deflection trailing edge airfoil, and it is more prominent at negative angle of attack. When the angle of attack is -4 °, the lift coefficient of the rigid-flexible airfoil is 1.66 times that of the traditional rigid airfoil, and when the angle of attack is 8 °, the lift coefficient of the rigid-flexible airfoil is 1.12 times that of the traditional rigid airfoil. The reason is that the pressure distribution of the rigid-flexible coupling deflection mode is smoother and more continuous at the trailing edge, and the area enclosed by it is always larger, so the lift enhancement effect is more obvious.

Fig. 8 shows the lift-drag ratio versus angle of attack for two different deflection modes when the trailing edge deflection angle is 30 °. It can be seen from Fig. 8 that under the same trailing edge deflection angle, the lift-drag ratio of the rigid-flexible hybrid deflection trailing edge airfoil is always larger than that of the rigid deflection trailing edge airfoil, but with the increase of the angle of attack, the lift-drag ratio advantage of the rigid-flexible hybrid deflection airfoil will be weakened. When the angle of attack is -4 °, the lift-drag ratio of the rigid-flexible airfoil is 2.86 times that of the traditional rigid airfoil, and when the angle of attack is 8 °, the lift-drag ratio of the rigid-flexible airfoil is 1.37 times that of the traditional rigid airfoil. Compared with the SM _ 3 configuration in reference [30], the airfoil is set to achieve flexible downbending deformation at 70% ~ 90% of the chord length. Through calculation, the lift characteristics and lift-drag ratio characteristics of the rigid-flexible mixed variable camber airfoil and the configuration show similar results.

In order to verify the advantages of the lift-drag characteristics of the rigid-flexible hybrid deflector airfoil more comprehensively, combined with the actual flight conditions of the transport aircraft, the lift per unit span of the wing in normal cruise is taken as the target.The deflection configurations to achieve the target lift under the low-speed cruise condition and the landing condition are optimized and calculated respectively, as shown in Table 2, and verified by CFD.

It can be seen from Table 2 that when the flight condition is low-speed cruise, the rigid-flexible hybrid airfoil does not need rigid deflection, but only drives the flexible section and the follow-up section, which account for 21% of the chord length of the airfoil, to reach the equivalent deflection angle of 8.7 ° to obtain the target lift, while the traditional rigid-deflection airfoil needs to deflect the overall trailing edge, which accounts for 30% of the chord length of the airfoil, by 7.3 ° to achieve the target lift. When the flight condition is the landing condition, which generally corresponds to the condition that the trailing edge of the wing is maximally bent down, the rigid-flexible hybrid airfoil only needs to be rigidly deflected by 3.78 degrees, flexibly deflected by 15 degrees, and the overall equivalent trailing edge bent down by 16.05 degrees to achieve the target lift, while the traditional rigid-deflected airfoil needs to be overall bent down by 29 degrees to achieve the target lift.

The above two airfoils are simulated and verified by CFD under different working conditions. When the flight condition is low-speed cruise and the same lift requirement is met, the overall trailing edge deflection angle of the traditional rigid deflection airfoil is 7.3 °, and its lift-drag ratio is 43.78, while the rigid-flexible deflection airfoil.The same effect can be achieved only by driving the flexible wing, which accounts for 21% of the chord length of the airfoil, downward by 8.7 °, and the lift-drag ratio is 45.53, which is 3.9% higher than that of the traditional rigid deflection, thus greatly improving the economy of transport aircraft^[31]. It can be seen that the rigid-flexible hybrid compliant variable trailing edge wing scheme provides a new design idea of high-lift device, and the compliant downbend with short chord can effectively improve the lift-drag ratio while maintaining the same high-lift effect^[32].

It can be seen from Fig. 9 and Fig. 10 that the two deflection modes have similar pressure distribution, velocity distribution and flow field characteristics under the low-speed cruise condition, and there is no serious flow separation at the tail of the airfoil due to the small angle of attack and the deflection angle of the trailing edge.

When the flight condition is landing and the target lift is achieved, the overall trailing edge deflection angle of the traditional rigid deflection airfoil is 29 °, and its lift-drag ratio is 9. 03, while the overall equivalent downward bending angle of the trailing edge of the airfoil corresponding to the rigid-flexible hybrid deflection is only 16. 05 °, and its lift-drag ratio is 16. 11, which is 78. 41% higher than that of the traditional rigid deflection. It can be seen that the deflection form greatly improves the continuity and smoothness of the airfoil trailing edge deflection, and makes it have better lift-drag characteristics. The advantage of lift-drag ratio is more obvious in large angle downbend.

Fig. 11 shows the pressure distribution nephogram corresponding to the two configurations in the landing state. By comparing Fig. 11 (a) and Fig. 10 (B), it can be seen that the area of the low pressure area formed by the rigid-flexible hybrid deflection airfoil at the tail is significantly smaller than that of the traditional rigid deflection airfoil. Fig. 12 shows the velocity nephogram and streamline corresponding to the two configurations under this condition. It can be seen that the flow separation starting position at the trailing edge of the rigid-flexible deflector airfoil in Fig. 12 (a) is at X/C = 0.8, and C is the chord length. The flow separation starting position at the trailing edge of the traditional rigid-deflector airfoil in Fig. 12 (B) is at X/C = 0.57. Comparing the two figures, the scale of the vortex structure formed by the flow separation at the trailing edge of the airfoil in Fig. 12 (B) is significantly larger than that of the airfoil in Fig. 12 (a). It can be seen that the traditional rigid deflection airfoil will cause premature flow separation, which seriously affects the aerodynamic efficiency, while the rigid-flexible hybrid deflection airfoil is more difficult to cause flow separation, and its separation point is closer to the trailing edge, so the aerodynamic efficiency is better. Fig. 13 shows the pressure distribution on the upper and lower surfaces of the airfoil corresponding to the two configurations under landing conditions. By comparing Fig. 13 (a) and Fig. 13 (B), it can be clearly seen that the rigid-flexible hybrid deflection airfoil has better continuity and smoothness at its deflection, and the mutation effect is smaller.

The aerodynamic design scheme of the optimal variable camber configuration is determined, and the important work of the variable trailing edge wing design is to realize the structure and deformation control design of the rigid-flexible hybrid wing^[33]. In this paper, the specific optimization design process of the rigid-flexible hybrid wing is omitted, and only the design principle and results are briefly described. The key to the design lies in the design of the compliant rib, which is designed by using the load path method for topology optimization, as shown in Figure 14. The load path method is similar to the basic structure optimization method in the initial layout, which connects the characteristic points in the design domain through the transfer form of the load path, and controls the diversity of the topological solution through the position and number of the intermediate points. Firstly, the appropriate number of nodes is arranged in the design domain, and the initial topological structure of the design domain is given by the full base configuration. Then, by parameterizing and discretizing the load path, different path forms from the load input point to the load output point are formed. Then, by using appropriate design variables, such as path number, path weight, etc., the optimal path from the load input point to the load output point under the restricted condition is obtained through the optimization algorithm combined with the finite element analysis, and finally all the structures covering these paths are combined to obtain the final topology optimization result.

In Fig. 15, point 1 is the input point of the driving load; Points 1 to 6 are output points on the upper and lower flange surfaces respectively; Points 7 ~ 11 are the intermediate nodes of the design domain; Point 12 is a fixed point. Firstly, the Dijkstra greedy algorithm in computer graph theory is used to solve the shortest path from the input point to each output point, and then the YEN algorithm is used to solve the first K load transfer paths from the input point to each output point through the shortest path. When K = 4, all paths have covered all initial base structures.

According to different paths, whether each beam structure in the design domain is called in a certain path is analyzed by the finite element. If it is called, it is retained in the finite element model. If it is not used, the beam is deleted from the finite element model. At the same time, the size and shape of the structure are optimized alternately, and the variable cross-section curved beam with better stability and deformability is introduced to replace the original uniform cross-section straight beam.Taking the fitting degree between the variable camber configuration of the compliant segment structure under the driving force and the previously obtained optimal variable camber configuration as an evaluation criterion, establishing the minimum quadratic variance between the actual variable camb configuration and the target configuration as an objective function of the space optimization process:^[34]

The optimal structural configuration meeting the material strength requirement under the target deformation requirement is obtained, as shown in Fig. 16.

In order to improve the load bearing capacity of the variable camber trailing edge, the real aerodynamic load is used as the technical index to further optimize the design of the variable camber trailing edge. The pressure coefficients of the upper and lower wing surfaces are shown in Fig. 17, taking the 15 ° downbend of the compliant section during landing as the analysis condition to calculate the true aerodynamic load per unit span. Through the analysis of the force transmission characteristics of the wing, it is necessary to arrange five ribs within the unit span, and to extend the thickness of a single rib to 20 mm, so as to achieve the required stiffness and structural strength for bearing. Formula (5) is the chordwise distribution of aerodynamic loads on the upper and lower edges of the rib:

Where: X is the position in the chordwise direction; Y is the corresponding aerodynamic load. Due to the increase of structural stiffness, the driving force of a single pneumatic muscle needs to be increased to 3250 N to achieve the desired structural deformation. In this paper, only the variable camber capability of the structure is verified, so the influence of aerodynamic load is not considered.

In the skin design, the rigid rotation section is designed with a pull-out skin structure according to the requirement of keeping the wing surface smooth and continuous during deformation. As shown in Fig. 18, the upper and lower wing skins extend into the leading edge of the wing box along the skin sliding tracks respectively, and are connected by springs fixed on the leading edge spar. When the rigid rotating section of the wing deflects downward, the upper wing skin will be pulled out along the track, and the tension spring is in a state of tension, and the skin will be tensioned by the restoring force of the spring.The lower wing skin is spring pre-tensioned in the initial position, and after being inserted into the wing box along the track, the spring remains in tension, providing a certain out-of-plane stiffness for the skin. The pull-out skin design keeps the wing structure relatively smooth and continuous at the leading and trailing edge skin spacing. At the flexible section, the upper and lower wing surfaces are currently designed with sliding skins, which are connected to the rib structure through tensioned sliding cables and slider guide rail devices to ensure the compliance of the connection, so that it can deform with the rib at the same time. Because the upper and lower skins are not sealed at the end of the rib, which may affect the aerodynamic efficiency and the strength of the skin structure, the motion trajectories of the upper and lower skins at the end of the rib are analyzed.The skin connection reinforcement structure was designed and manufactured by 3D printing technology, as shown in Fig. 19, which was fixed between the upper and lower skins, so that the upper and lower skins at the end of the rib were always closely attached during the deformation process.

The optimized wing rib structure is processed and manufactured, and the selected material is 7075-T6 aviation hard aluminum alloy. The rigid rotation section is driven by a servo motor and a crank-connecting rod mechanism, and the horizontal motion of the actuation output end is converted into the fixed-axis rotation of the wing rib structure.Each rib in the flexible deformation section is driven by two pneumatic tendons, and the air intake of the pneumatic tendons is controlled by the voltage, so as to change the internal pressure, and finally the axial deformation of the pneumatic tendons is converted into driving force output^[35-36]. After assembly, a 3-rib wing-box model is obtained, as shown in Figure 20.

The deformation verification test of the trailing edge of the designed morphing wing was carried out at a maximum equivalent deflection angle of 30 °. The three-rib wing box structure is first rigidly deflected under the drive of the servo motor, and the rotation angle and the horizontal actuation distance of the driver can be directly calculated from their geometric relationship. When the rigid rotation section is bent downward by 20 °, the horizontal displacement driven by the servo motor is 21 mm. When the rigid rotating section rotates to the specified position, the servo motor driver is locked. On this basis, the pneumatic tendon was used to drive the flexible segment to bend downward while maintaining the rigid rotation. The test results show that the equivalent deflection angle of the compliant segment structure can reach 15 ° and the overall trailing edge can reach 30 ° when the pneumatic muscle is applied with a pressure of 3.9 bar (6.5 V pneumatic valve driving voltage), as shown in Figure 21.

In this paper, a design scheme of rigid-flexible hybrid deflection airfoil is proposed by considering the factors of variable amplitude, deformation continuity, deformation stress and driving efficiency, and the aerodynamic efficiency is compared with that of the traditional deflection mode.Then the rib structure and actuator layout scheme are given according to the actual aerodynamic load. Finally, the deformation capacity of the rigid-flexible hybrid variable camber trailing edge wing structure and skin is verified by the test.

Firstly, the parametric geometric description method of wing variable camber configuration based on the airfoil mean camber line was established.Corresponding to the deflection target of 30 ° of the trailing edge, the lift-drag ratio of the airfoil was used as the evaluation function, and finally the optimal downward deflection angles of the rigid and flexible parts and the variable camber curve of the flexible part were obtained by optimization, so as to obtain the optimal variable camber configuration.

Secondly, CFD simulation is used to compare the aerodynamic characteristics of the two airfoils at different angles of attack, such as lift coefficient and lift-drag ratio. The results show that the lift coefficient and lift-drag ratio of the rigid-flexible airfoil are better than those of the traditional rigid-deflected airfoil at angles of attack from-4 ° to 8 °, and the effect is more obvious at small angles of attack, because the flow separation at the tail leads to the increase of drag and the decrease of aerodynamic efficiency when the angle of attack increases.

Then, compared with the actual flight conditions of the aircraft, the optimal downward deflection angles of the two deflection airfoils at low speed cruise and landing conditions are calculated respectively, with the wing lift per unit span as the optimization objective. In order to provide the target lift at low speed cruise, the rigid-flexible hybrid wing only needs to drive the flexible section down 8. 7 ° to meet the requirement, while the traditional rigid wing needs to drive the whole trailing edge section down 7. 3 °. In the landing condition, the rigid-flexible hybrid wing needs to deflect the rigid section by 3. 78 °, the flexible section by 15 °, and the overall equivalent deflection by 16. 1 ° to achieve the optimization goal, while the traditional rigid wing needs to deflect the entire trailing edge section by 29 ° to meet the optimization goal. Furthermore, the flow field characteristics are calculated by CFD simulation, and it is found that the rigid-flexible hybrid deflection airfoil has greater advantages in the case of large deflection angle, and its lift and drag are increased by 78% compared with the traditional rigid deflection airfoil in the landing condition. The flow separation point is located closer to the trailing edge of the airfoil, only at X/C = 0.8, and the vortex structure at the trailing edge is smaller, resulting in higher aerodynamic efficiency.

The chordwise pressure distribution of the upper and lower wing surfaces is given by the actual aerodynamic load calculation, and the spanwise thickness of the ribs and the number of ribs per unit span are determined by the finite element analysis. Because the weight of a single compliant rib structure is only 0. 35 kg, and the mass of the pneumatic muscle and the control module is much lower than that of the traditional hydraulic drive device, it also has great advantages in structural weight reduction.

Finally, the principle and method of structural optimization of the compliant rib and the skin design method of the rigid rotating section and the compliant section are introduced, and the model manufacturing and the deformation capacity test without considering the aerodynamic load are carried out.The rigid segment and the flexible segment at the trailing edge of the rigid-flexible hybrid deflection were driven by the servo motor and the pneumatic tendon, respectively.After the rigid segment was bent to 20 °, when the driving voltage of the pneumatic tendon reached 6. 5 V, the rigid-flexible hybrid deflection angle reached the design value of 30 °, and the deformation state under different voltages was recorded.