Delaying stall of morphing wing by periodic trailing-edge deflection

CJA 1423 16 October 2019 Chinese Journal of Aeronautics, (2019), xxx(xx): xxx–xxx

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Chinese Society of Aeronautics and Astronautics & Beihang University

Chinese Journal of Aeronautics [email protected] www.sciencedirect.com

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Delaying stall of morphing wing by periodic trailing-edge deflection

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Zi KAN, Daochun LI *, Jinwu XIANG, Chunxiao CHENG

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School of Aeronautic Science and Engineering, Beihang University, Beijing 100083, China

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Received 8 October 2018; revised 11 September 2019; accepted 11 September 2019

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KEYWORDS

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Aerodynamics; Computational fluid dynamics; Flexible structures; Morphing wings; Stall angle

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Abstract Morphing wings can improve aircraft performance during different flight phases. Recently research has focused on steady aerodynamic characteristics of the morphing wing with a flexible trailing-edge, and the unsteady aerodynamic and stall characteristics in the deflection process of the morphing wing are worthy further investigation. The effects of the angle of attack and deflection rate on aerodynamic characteristics were examined, and based on the aerodynamic characteristics of the morphing wing, a method was developed to delay stall by using the flexible periodic trailing-edge deflection. The numerical results show that the lift coefficients in the deflection process are smaller than those in the static situation at small angles of attack, and that the higher the deflection rate is, the smaller the lift coefficients will be. On the contrary, at large angles of attack, the lift coefficients are higher than those in the static case, and they become larger with the increase of the deflection rate. Further, the periodic deflection of the flexible trailing-edge with a small deflection amplitude and high deflection rate can increase lift coefficients at the critical stall angle. Ó 2019 Production and hosting by Elsevier Ltd. on behalf of Chinese Society of Aeronautics and Astronautics. This is an open access article under the CC BY-NC-ND license (http://creativecommons.org/ licenses/by-nc-nd/4.0/).

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1. Introduction

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Morphing wing aircraft can adapt its airfoil shapes to suit the mission conditions with high operating efficiency and better aerodynamic performance. The recent government-sponsored activities, historical perspectives and future challenges of mor-

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* Corresponding author. E-mail address: [email protected] (D. LI). Peer review under responsibility of Editorial Committee of CJA.

Production and hosting by Elsevier

phing aircraft systems,1 and the classification, performance benefits, and enabling technologies of the morphing wing2 have been discussed. Recently, Li et al.3 provided an overview about the most prominent examples of morphing wings, as well as the methods and tools commonly deployed for the design and analysis of these morphing wing concepts. The flexible trailing-edge has become the most commonlyused method for the morphing wings design due to its significant benefits and convenient implementation.4–7 A wind tunnel test was carried out to prove that the morphing wing with flexible trailing-edge had better aerodynamic characteristics than a traditional airfoil.8 Li et al.9 studied the aeroelastic response of a morphing flap showing that the morphing of airfoil can reduce critical flutter speed. A flexible trailing-edge with

https://doi.org/10.1016/j.cja.2019.09.028 1000-9361 Ó 2019 Production and hosting by Elsevier Ltd. on behalf of Chinese Society of Aeronautics and Astronautics. This is an open access article under the CC BY-NC-ND license (http://creativecommons.org/licenses/by-nc-nd/4.0/). Please cite this article in press as: KAN Z et al. Delaying stall of morphing wing by periodic trailing-edge deflection, Chin J Aeronaut (2019), https://doi.org/10.1016/j. cja.2019.09.028

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deformable ribs10 and a morphing wing with a double corrugated variable camber11 were designed to improve aerodynamic performance. A flexible variable camber trailing-edge flap was designed by Lu et al.12 to improve flight efficiency during takeoff, cruise and landing states. Moreover, supported by governments, the American company FlexSys designed a mission adaptive compliant wing of high altitude endurance aircraft with a smooth and variable trailing-edge, and the test results showed that the lift-to-drag ratio could be increased by 3.3% under cruise conditions.13,14 Boeing and NASA jointly designed a Variable Camber Continuous Trailing-edge Flap system which was found to potentially reduce nearly 10% drag in the cruise condition.15 European Union supported the Smart Intelligent Aircraft Structures (SARISTU) project to accomplish the design of the morphing wing system and manufacture flexible morphing structures.16 Morphing wings are used not only in fixed wing aircraft, but also in other devices and vehicles, such as wind turbine17,18 and rotor craft.19–22 Aerodynamic analysis plays an important role in the process of translating morphing concepts into designs. Due to the high computational efficiency, the extension of Weissinger’s method23 and the vortex lattice method24 are often used to obtain steady aerodynamics of the morphing wing. Considering the flexibility of morphing wings, the unsteady vortex lattice method,25,26 the doublet lattice method27,28 and the CFD method are adopted to estimate the unsteady aerodynamics. By solving the Reynolds-Averaged Navier-Stokes (RANS) equations, the flow transition over the morphing wing surface was predicted, and the numerical results were found in good agreement with the wind tunnel experimental results for the pressure distribution.29.The wind tunnel experimental data and numerical simulation results were compared to prove that the CFD method could be successfully used to perform optimization of the morphing wing.30 The CFD method exhibits an adequate application in the cases with large angles of attack and strong viscous flow. Some studies have researched the aerodynamic and stall characteristics of morphing wings with CFD method. Lyu and Martins31 adopted the RANS equations to study adaptive morphing trailing-edge wings, observing 1% drag reduction in on-design conditions, and 5% drag reduction near off-design conditions. Abdessemed et al.32 used the CFD method to analyze a NACA0012 airfoil with a morphing trailing-edge; the results revealed that the lift-drag ratio increased by 6.5% and that the local flow field was influenced by the morphing motion. The stall and aerodynamic performance of a NACA0012 airfoil fitted with flexible trailing-edge was investigated by RANS numerical simulations, showing that the morphing surface could delay the onset of flow separation to achieve optimal aerodynamic performance.33,34 In addition, Wolff et al.17,35 solved the RANS equations to investigate the aerodynamic and stall characteristics of a deformable trailing-edge of wind turbine blade airfoil. The results showed that while deflecting the trailing-edge at angles of attack near stall, the transient lift is higher than the steady lift. Ahaus investigated the dynamic stall behavior of a pitching airfoil with a morphing trailing-edge of helicopters.20–22 In this study, a method for delaying stall by periodic trailing-edge deflection is proposed based on the unsteady aerodynamics of a morphing wing in the deflection process using the CFD method with dynamic mesh.

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A flexible trailing-edge can achieve a chord-wise and span-wise differential camber variation with the same structural system, providing a smooth contour with no additional gap. In this research, the NACA0012 airfoil with a flexible trailing-edge was modeled for the numerical simulation of aerodynamics. To study the aerodynamic characteristics of the morphing wing, the CFD method with dynamic mesh and UserDefined Functions (UDFs) were adopted.

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2.1. Model

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A model based on the NACA0012 airfoil with a flexible trailing-edge was designed to numerically simulate the aerodynamics of morphing wings. A morphing wing and a conventional wing with a flap are shown in Fig. 1. The morphing section of the morphing wing is in the 60%-100% chord of the airfoil, and the fixed section is in the 0–60% chord of the airfoil. The angle between the chord line and the line connecting the deflection axis to the trailing-edge is the morphing trailing-edge at the deflection angle b which varies in a range of ±20°, with the downward direction as the positive direction. V1 represents the free-stream velocity, and the angle between the chord line and the direction of V1 is the angle of attack a.

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2.2. Modeling method and morphing method

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Based on Refs. 36,37, this study adopts the parabolic form of morphing. Given its constant thickness, the morphing wing with a flexible trailing-edge can be modeled according to the

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Fig. 1 wing.

Comparison between morphing wing and conventional

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distribution of the airfoil thickness as long as the morphing form of the mean camber line is known.

chord is unchanged. The equation of the mean camber line of the morphing section is

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2.2.1. Modeling method

y ¼ tan bðtÞ=fðbðtÞÞ  x

ð3Þ

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As shown in Fig. 2, the initial deflection point is set as the origin of coordinates (0, 0), and coordinate value of the morphing trailing-edge endpoint is (l, Dz). According to the geometric relationship in this figure, the deflection angle and coordinate value of the trailing-edge endpoint can be expressed as tanb = Dz/l. Then the trajectory equation of the mean camber line is

and the position of grid nodes of the morphing wing (x(t), y(t)) can be obtained at each moment. Finally, the trailing-edge can continuously deflect with the updating position of grid nodes. It should be noted that during the continuous deflection of the trailing-edge, the a remains unchanged.

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2.3. Numerical method and validation

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In the numerical calculation, the unsteady incompressible RANS equations with the Spalart-Allmaras turbulence model are discretized with the finite volume method. A second-order upwind scheme, a second-order central difference scheme, and the SIMPLE algorithm are used for the convection terms, diffusion terms, and pressure-velocity coupling, respectively. The computational domain sketch and unstructured mesh layout are shown in Fig. 3. The radius of the computational domain is 30 times the chord length c of the airfoil, the boundary of the computational domain is set as a pressure-far-field, and the airfoil is set as a no-slip static wall. As shown in Fig. 3(b), the grid is refined near the airfoil for more accurate results. To validate the accuracy of the numerical simulation, this study compared the lift and moment coefficients of the airfoil NACA0012, using the experimental data and results of the extended-ONERA method.38,39 Although the model does not have a flexible trailing-edge, this airfoil experienced the harmonic pitching motion about the quarter-chord. The form of the AoA is

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y ¼ tanb=l  x

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It is known that the length of the mean camber line of the airfoil is 0.4 m, remaining unchanged during the morphing R l qffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffi process. By integrating the mean camber line 0 1 þ ðy0 Þ2 dx, the following equation can be obtained: ! pffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffi pffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffi l 1 1 þ 4 tan2 b  2 tan b 4 1 þ 4 tan2 b  ln pffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffi ¼ 0:4 8 tan b 1 þ 4 tan2 b þ 2 tan b ð2Þ

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According to Eqs. (1) and (2), the relationship between deflection angle and coordinate value of the morphing trailing-edge endpoint is obtained. If the deflection angle is determined, (l, Dz) can be determined, and then the equation of the mean camber line can be obtained. According to the thickness distribution of the NACA0012 airfoil, the morphing airfoil shape is obtained on the basis of the mean camber line trajectory. The equations of the upper and lower surfaces of the morphing wing can also be obtained. Then the morphing wing with a flexible trailing-edge is modeled.

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2.2.2. Morphing method

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In this research, the dynamic deflection of a flexible trailingedge needs to be simulated to study the unsteady aerodynamic characteristics of the morphing wing. The user-defined functions were adopted to achieve the morphing of the flexible trailing-edge by controlling the position of grid nodes. For example, the deflection angle varies in the form b(t) = b0  sin(2pft) with f representing the deflection rate (cycles per second) and b0 the maximum deflection angle of the trailingedge. In the deflection process, the region of 0–60% airfoil

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Fig. 2

Trajectory of mean camber line.





aðtÞ ¼ 9:97 þ 9:88  sinð2pftÞ

ð4Þ

with the Mach number, Reynolds number, and deflection rate f being 0.302, 3.78  106 and 8.459 Hz, respectively. The com-

Fig. 3

Computational domain and unstructured mesh layout.

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parison between the lift and moment coefficients of the airfoil verified the accuracy of the CFD method with dynamic mesh. Fig. 4 shows that the lift coefficients CL obtained from both the CFD method and extended-ONERA method are slightly different from the experimental data, but trend along similar lines, especially during the increase of the a. A comparison of the moment coefficients Cm of the CFD, extendedONERA method and experimental data, shows some differences. However, the moment coefficients of the CFD results are closer to those of the experimental data, compared with those of the extended-ONERA method. Therefore, the CFD method is sufficiently accurate in simulating the unsteady aerodynamics.

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3. Results and discussion

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In this section, the transient aerodynamic characteristics of the morphing wing with a flexible trailing-edge in the deflection process are provided, and then a method for delaying stall of the morphing wing is introduced.

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Z. KAN et al. tion angle 20°, and then remains constant. Meanwhile, the a is fixed in the deflection process. The Mach number is 0.1 and Reynolds number is 2.3  106. The deflection angle varies in the form b(t) = b0  sin(2pft). 3.1.1. Effects of a and deflection rate on unsteady aerodynamic characteristics

The aerodynamic characteristics of the morphing wing in the deflection process were studied. The trailing-edge of the wing begins to deflect downward at the fourth second until the angle of the morphing trailing-edge increases to the maximum deflec-

The effects of the a and deflection rate on the aerodynamics of the morphing wing are discussed as below. The lift coefficients of the static and dynamic cases are shown in Fig. 5. The static case means that the trailing-edge deflection angle is always b0, and the dynamic one means that the initial deflection angle is 0°, and finally remains being b0 after the downward deflection. In Fig. 5, for both b0 = 20° and 15°, the transient lift coefficients are smaller than those of the static case at small a. Higher deflection rate leads to smaller transient lift coefficients. On the contrary, the transient coefficients are higher than those of the static case at large a; transient lift coefficients are larger with the increase of deflection rate. Furthermore, the transient lift coefficients decrease slowly in the stall stage, whereas lift coefficients of the static case drop rapidly. Therefore, the dynamic cases have wider range of effective a. To illustrate the effects of the a and deflection rate in detail, the transient lift coefficients are extracted in the deflection process. Fig. 6 shows the curves of the lift coefficients differences (DCL = CL transient  CL static) versus time for a = 4° and 12° at three different deflection rates. In Fig. 6(a), the trailingedge deflection rate is 1 Hz, the deflection angle varies in the

Fig. 4 Comparison among CFD results, experimental data and extended-ONERA method.

Fig. 5 rates.

3.1. Unsteady aerodynamic characteristics of morphing wing in deflection

Static and transient lift coefficients at different deflection

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5 1.67 Hz, respectively. Obviously the higher the deflection rate is, the larger the transient lift coefficients will be. On the contrary, for a = 4°, the transient lift coefficient at b0 = 20° decreases by 4.6%, 7.7% and 9.3% when the deflection rates are 1 Hz, 1.33 Hz and 1.67 Hz, respectively. Therefore, at large a, especially at the critical stall angle, the downward flexible trailing-edge deflection improves unsteady aerodynamic characteristics of morphing wings.

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3.1.2. Flow field of morphing wing in deflection process

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To explain the results in Section 3.1.1, a more detailed study about flow separation and pressure changes was conducted. The a of the static and dynamic cases are both 12°, the maximum deflection angle is 20°, and the deflection rate is 1.33 Hz. As shown in Fig. 7, compared with the dynamic cases, the static case shows a larger vortex attached to the upper surface of the trailing-edge, which indicates a larger flow separation. Thus, the lift coefficients of the static case are lower than those of the dynamic one. Fig. 7(b) and (c) display the streamlines and pressure contours when the deflection angles increase to 19.3° and 20°, respectively. When the separation near the trailing-edge occurs and begins to expand forward, the trailing-edge has already reached another position due to the continuous morphing. Thus, the evolution of the flow field always lags behind the trailing-edge morphing. In addition, the separation in Fig. 7(b) is smaller than that in Fig. 7 (c), which can explain the phenomenon in Fig. 6 that the transient lift coefficient increases up to the peak and then decreases to a stable value. These results show that, during the deflection, the

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Fig. 6 Differences between transient and static lift coefficients versus time for a = 4° and 12° at different deflection rates.

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form b(t) = 20°  sin(2pft), and the static lift coefficient is 2.15 at a = 4°, and 2.09 at a = 12°. As the trailing-edge reaches the maximum angle 20° at t = 4.25 s, DCL is 0.073 at a = 12°, and 0.1 at a = 4°. In addition, there is a peak difference 0.131 at a = 12° and t = 4.192 s. In Fig. 6(b), the trailing-edge deflection rate is 1.33 Hz. The deflection angle varies in the form b(t) = 20°  sin(2  1.33pt). As b is 20° at t = 4.188 s, the transient difference DCL is 0.143 at a = 12°, and DCL is 0.166 at a = 4°. The peak difference is 0.186 at a = 12°. In Fig. 6(c), the deflection rate is 1.67 Hz. The deflection angle varies in the form b(t) = 20°  sin(2  1.67pt). As b reaches 20° at t = 4.15 s, the transient difference DCL is 0.163 at a = 12°, and DCL is 0.16 at a = 4°. The peak difference is 0.21 at a = 12°. Therefore, the transient lift coefficient first increases up to the peak and then decreases slightly to a stable value during the process of the downward trailing-edge deflection, and the difference of the lift coefficient is always larger than 0 at a = 12°, and smaller than 0 at a = 4°. Furthermore, compared with the static lift coefficients with a = 12°, the transient lift coefficient at b0 = 20° increases by 3.5%, 6.8% and 7.8% when the deflection rates are 1 Hz, 1.33 Hz and

Fig. 7 Flow field and pressure contours of static and dynamic cases at f = 1.33 Hz, and a = 12°.

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morphing wing has better stall and aerodynamic characteristics at the critical stall angle.

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The analysis so far has shown that the morphing wing with a flexible trailing-edge has a larger lift coefficient than that of the static case at large a in the deflection process. Thus, we conducted further research to apply this property. This subsection introduces a method for improving stall characteristics with periodic flexible trailing-edge deflection. In the trailing-edge deflection process, the a remains constant. In this part, this deflection is in the form b(t) = 20° + b1  sin(2pft); the static case means that f = 0, and the trailing-edge deflection angle is fixed as 15°. The dynamic case means that the flexible trailing-edge deflects periodically with 15° as an equilibrium position. The mean lift coefficients of those cases are shown in Fig. 8. It can be seen that the critical stall angles of the static case are about 13°, and in the dynamic cases, the lift coefficient has still no obvious change when the a reaches to 16°. The mean dynamic lift coefficient of the periodic trailing-edge deflection is larger than that of the static case in the stall stage, and the stall angle increases. Then, a more detailed study was conducted, as shown in Fig. 8 (b). The lift coefficients at a = 14° and 16° correspond to the left and right Y-axes, respectively. The result shows that the smaller the deflection amplitude is and the higher the deflection rate is, the larger the dynamic mean lift coefficient will be. Thus, the trailing-edge periodic deflection with a small deflection

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Fig. 8

Static and mean dynamic lift coefficients.

amplitude and high deflection rate can improve the performance of morphing wings in the critical stall stage. The case of a = 16° will be used as an example to analyze the effect of the periodic trailing-edge deflection on lift coefficients. As shown in Fig. 9(a), the static lift coefficient begins to decrease when the deflection angle is 10°, while the lift coefficient of the periodic trailing-edge morphing is still increasing. The dynamic lift coefficients changing with the deflection angle present a hysteresis loop, which moves in a clockwise direction and can be seen in Fig. 9(b). Furthermore, the higher the deflection rate is, the more sharply the hysteresis loop slopes; the smaller the deflection amplitude is, the narrower the hysteresis loop is. This phenomenon can be briefly explained in terms of flow separation. The flow of the static case begins to separate at the deflection angle of 10°, so the lift coefficients are smaller than those of the dynamic cases. For the dynamic cases, during the downward process of the periodic deflection, the higher the deflection rate is, the more slowly the separation expands, leading to a large lift coefficient at the maximum deflection angle. However, during the upward process of the periodic deflection, the reattachment is more difficult with high deflection rate, resulting in a small lift coefficient at the minimum deflection angle. So a sharp slope of the hysteresis loop is generated. In addition, the small deflection amplitude causes a more insignificant difference of the lift coefficients during periodic deflection of the trailing-edge, and a narrower hysteresis loop.

Fig. 9 Static and periodic lift coefficients of trailing-edge deflection at a = 16°.

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4. Conclusions The aerodynamic characteristics of the morphing wing with a flexible trailing-edge are investigated numerically in this study. The unsteady aerodynamics in trailing-edge deflection process with different a and deflection rates are investigated. A method is proposed to delay the stall of the morphing wing by periodic trailing-edge deflection. (1) During the downward deflection of the flexible trailingedge, the transient lift coefficients are smaller than those of the static case at small a. Higher deflection rate leads to smaller transient lift coefficients. However, at large a, especially in the critical stall stage, the transient lift coefficients are larger than those of the static case, and higher deflection rate results in larger transient lift coefficients. (2) By periodic deflection of the trailing-edge, the airfoil stall can be delayed. The results show that the periodic deflection of the trailing-edge can increase the lift coefficients. Furthermore, a higher lift coefficient will be obtained with a small deflection amplitude and high deflection rate. The lift coefficients present a hysteresis loop, moving in a clockwise direction.

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Acknowledgments

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The authors are grateful to the anonymous reviewers for their critical and constructive review of the manuscript. This study was supported by the National Natural Science Foundation of China (Nos. 11402014 and 11572023).

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Please cite this article in press as: KAN Z et al. Delaying stall of morphing wing by periodic trailing-edge deflection, Chin J Aeronaut (2019), https://doi.org/10.1016/j. cja.2019.09.028