# Combined droop nose and trailing-edges morphing effects on airfoils aerodynamics

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## Abstract

Airfoils’ morphing offers benefits to traditional aerodynamics characteristics flight such drag saving and better maneuver skills for aircrafts by different means. The current study introduces a parametric optimization for the effect of droop nose and trailing-edges morphing on airfoils aerodynamics characteristics. The study has conducted using the X-FOIL imported in MATLAB program and ANSYS Fluent software. The first used as a prelude parametric study for the numerical simulations. The morphed airfoil configuration parameterized with four variables: leading and trailing edges deflections, *δ*_{1}, *δ*_{2} and the morphing lengths *E*_{1}, *E*_{2}. The results from X-FOIL showed increase in stall angle up to 26° and lift to drag ratio up to 120. The more accurate simulation results showed increase of stall angle up to 20° and lift to drag ratio increased up to 6.22% compared to basic airfoil at separate morphed shapes. The lift coefficient increased also up to 1.34. The study also introduces morphing control to insure maximum lift coefficient during the variety of operating angle of attack.

## Keywords

Morphing Droop nose Trailing-edges Experimental measurement Aerodynamic control## 1 Introduction

The high-lifting devices in air craft such flaps, slats and ailerons represent overweight and extra complexity. This extra weight of the mechanism, complexity of the operating and control systems is a problem from the design and operation point of view. By itself, the overall system complexity and structure weight are considerably increased. Unlike conventional wing control surfaces, morphing leading-edge and trailing-edge usually use the conformal structural deformation achieved through bending and twisting of structures to adaptively change wing shape, leading to potentially systems and reduced weight. Also, morphing allows a smooth change for geometry surface to outfit a certain aero load which is desirable for noise reduction.

Recent studies focused in such helpful method to improve the aerodynamic performance of different applications such wind energy, aircraft parts and component and etc. Yue et al. [1] have worked on aircraft maneuvering enhancement using morphing wing by increasing the roll angular velocity. They have used the approximation of quasi steady aerodynamics. Kamliya et al. [2] have showed that morphing with higher camber flap increase the lift coefficient while a little decrease the lift to drag ratio is observed. Gamble et al. [3] have used morphing hinge at airfoil trailing edge which was capable of delay the inception of stall to higher angle of attack but with slight reduction in lift coefficient. Kimaru and Bouferrouk [4] have showed experimentally the wing with morphing camber which has a better aerodynamic performance for different angles of attack than traditional wings.

Aerodynamic performance enhancement and structural mass reduction have been introduced using morphing skill applied to airfoil camper Vale1 et al. [5]. Jodin et al. [6] have showed that vibrating the airfoil trailing edge with high frequency and little amplitude was capable of rising the lift and reducing the drag consequently. Vasista et al [7] have studied the wing tip droop-nose using morphing technique and explained experimentally the visions from tests and gave some recommendations also they were not envisioned to improve the wing aerodynamic loads coefficients. As they mentioned, their study has lacked of aerodynamic studies for the problem to determine the effect of droop-nose on the flying performance.

A saving in fuel consumption have been achieved using morphing wingtips and growth in yaw power using twisting morphing and bending morphing Afonso et al. [8]. Morphing wing applications have extended to growth the photo voltaic collected energy during small angle of sun rise by correcting dihedral angle of small wing part near tip with acceptable variant in aerodynamics loads Wu et al. [9]. Many studies have focused on nonlinearity of the morphing wings. Zhang et al. [10] have introduced dynamic analyses for deploying wings and aerodynamic improvement using the thin airfoil theory. Hu et al. [11] have presented some aero-elasticity features and responses of morphing wings.

The current study introduces the aerodynamic characteristics of morphing airfoil including droop nose and trailing edge morphing. The parametric optimization is performed first using X-FOIL-MATLAB followed by numerical simulation study with ANSYS Fluent software. The studies aims to improve aerodynamics characteristics flight such drag saving and peter maneuver skills for aircrafts by different means such delaying stall inception and enhance the aerodynamics coefficients compared to the original airfoil.

## 2 Parametric study for the variation

The predication for the aerodynamic characteristics are calculated using X-FOIL imported to MATLAB; the X-FOIL applies the vortex lattice method (VLM) in order to solve for the aerodynamic characteristics for a given airfoil. The VLM gives a good prediction for the aerodynamic performance and to indicate the trend of some variation on a certain airfoil. The reason to use the VLM is that the high computational cost of using the CFD, the prediction resulted from the VLM could narrow the working region in order decrease the number of runs to be done using the CFD. The accuracy X-FOIL in comparison with CFD and experimental measurements has presented by Günel et al. [12].

*α*

_{STALL}. A parametric study is then conducted including sufficient number for different morphing geometries. A MATLAB script has been prepared to generate the morphed shape as function of four control parameters. The airfoil morphed profile is shaped using a deflection having a parabolic profile with three coefficients in which are calculated based on the

*E*

_{1},

*E*

_{2},

*δ*

_{1}and,

*δ*

_{2}values and the unmorphed profile of the airfoil. For the deflection which is denoted as δ for the leading edge:

*l*

_{1},

*l*

_{2}and

*l*

_{3}are constant to be determined by applying conditions at leading edge.

*c*is the airoil chord length

*t*

_{1},

*t*

_{2}and

*t*

_{3}are constant to be determined by applying conditions at trailing edge. Finally, considering the ranges discussed above the new profile becomes,

The current study intended on low speed aircraft up to 45 m/s flight speed and 1 km flight altitude where the Reynolds number change did not exceed 5% for a unity chord length. For that the Reynolds number is considered constant during the different flight regimes. The study includes high lift generation flight regime during takeoff and maneuvering where the critical parameter is the stall angle of attack. Also, the cruise flight regime where the maximum lift coefficient and minimum drag coefficient are most preferred.

^{6}, the density

*ρ*= 1.225 kg/m

^{3}, the inlet velocity

*V*= 43.822 m/s.

Morphing parameters

Parameter | Set 1 | Set 2 | ||||
---|---|---|---|---|---|---|

Minimum | Maximum | Step | Minimum | Maximum | Step | |

| − 0.03 | 0 | 0.015 | − 0.06 | − 0.045 | 0.0075 |

| 0.2 | 0.4 | 0.1 | 0.2 | 0.4 | 0.1 |

| 0 | 0.06 | 0.012 | 0 | 0.06 | 0.012 |

| 0.2 | 0.4 | 0.04 | 0.2 | 0.4 | 0.04 |

*δ*

_{2}and the morphing length

*E*

_{2}, at a different leading edge deflection

*δ*

_{1}and morphing length

*E*

_{1}. It is obvious that the trailing edge upward deflections mainly affects the stall angle of attack positively Fig. 2a, b, but as a cost for this, the \( C_{L} /C_{{D_{max} }} \), decreases dramatically in which can be compensated using the downward leading edge deflection as shown in Fig. 3a, b.

**E**_{2}is that the longer

**E**_{2}the lower stall angle of attack while it gives higher lift to drag ratio. This may be explained as the study combined effect of both

**δ**_{2}and

**E**_{2}, at high angle of attack the separation appears near trailing edge. The deflected trailing edge works on better attachment for the flow at high angle of attack. The effect of a pure trailing edge deflection can be represented by the Table 2. The stall angle of attack didn’t face any attenuation but it increased. The remarkable point is that the \( \left( {C_{L} /C_{D} } \right)_{MAX} \) has increased a lot more than that in case of a basic airfoil, the effect of increase of

**E**_{1}is that it mainly decreases the \( \left( {C_{L} /C_{D} } \right)_{MAX} \), and has no remarkable effect on the stall angle of attack. The leading edge deflection effect can be shown by Table 3

Effect of trailing edge morphing length

Airfoil | | | | \( C_{L} /C_{{D_{max} }} \) |
---|---|---|---|---|

Morphed1 | 0.06 | 0.2 | 20.62 | 45.29 |

Morphed2 | 0.06 | 0.4 | 19.74 | 62.23 |

Original | 0 | – | 17 | 96 |

Effect of leading edge morphing length

Airfoil | | | | \( C_{L} /C_{{D_{max} }} \) |
---|---|---|---|---|

Morphed1 | 0.03 | 0.2 | 18.11 | 140 |

Morphed2 | − 0.03 | 0.4 | 18.86 | 116 |

Original | 0 | – | 17 | 96 |

Effect of leading and trailing edges morphing length

| | | | | | \( C_{L} /C_{{D_{max} }} \) |
---|---|---|---|---|---|---|

1 | − 0.06 | 0.2 | 0.06 | 0.2 | 20.83 | 92.73 |

2 | − 0.06 | 0.2 | 0.06 | 0.4 | 20.83 | 130.4 |

3 | − 0.06 | 0.4 | 0.06 | 0.2 | 25 | 70.31 |

4 | − 0.06 | 0.4 | 0.06 | 0.4 | 25 | 93.95 |

## 3 Numerical simulations

The numerical simulations have conducted using ANSYS Fluent for Newtonian flow. The software solves 2D Navier–Stokes equations, incompressible, steady-state fluid flow using finite volume technique. The simulations have performed using workstation with the following characteristics: Intel core i7, 5th generation, 3 GHz processor (cash 4 MB) and 16 GB RAM. The geometry has created by ANSYS design modeler and the mesh generation has conducted using ANSYS mesh modeler using unstructured mesh with some modifications around the airfoil to capture the boundary layer and flow separation.

## 4 Grid sensitivity and verification

The sensitivity analysis had conducted to increase the accuracy of the grid with the results (lift coefficient, Drag coefficient, Y+ , etc.). The *Y*^{+} is a good indication of grid quality. For low-Reynolds number turbulence models the Y+ value ≤ 1.0 Ariff et al. [13]. The grid needs to be clustered near to wall to capture the boundary layer on the airfoil at different angle of attacks.

*Re*= 3 × 10

^{6}with standard air properties at sea level have considered where the density ρ = 1.225 kg/m

^{3}at T = 288 k temperature and dynamic viscosity μ = 1.7894 × 10

^{−5}kg/ms. The airfoil chord was c = 1 cm, the inlet velocity V = 43 m/s The domain of the computational model is a half circle along its length with radius 7

*c*extended with a square shape with 14

*c*length. The two parallel side to the chord and the circle half is flow inlet velocity boundary (Drichlet value) and right side is pressure flow outlet (Gauge pressure) as shown in Fig. 4. The grid is unstructured consists of two regions, where the inner region is a circle with diameter equal to 3C. The mesh around the airfoil is cluster to capture the boundary layer Fig. 5. Turbulent model selected according to recommendations illustrated in Aziz et al. [14].

^{5}cells. Figure 7 presents the percentage in mass flow errors between inlet and outlet boundaries for the different angle of attack to insure the grid consistency.

The mesh shape controlled parameters

Computational grids number | 1 | 2 | 3 | 4 | 5 | 6 | 7 | 8 | 9 | 10 |
---|---|---|---|---|---|---|---|---|---|---|

Max face size × 10 | 40 | 20 | 10 | 9.5 | 9 | 8.5 | 8.5 | 8 | 7.5 | 5 |

Near wall patch thickness × 10 | 1.2 | 5 | 7.5 | 7.5 | 1 | 1 | 1 | 1 | 1 | 1 |

Total number of cells × 10 | 1.2 | 1.4 | 2.6 | 2.7 | 2.9 | 3.1 | 3.3 | 3.4 | 3.7 | 7.1 |

## 5 Software validation with experiment

*Re*= 3 × 10

^{6}with standard air properties at sea level have considered. The airfoil chord was c = 20 cm, with the same inlet velocity considered in the experiment. The Spalar–Allmaras model has used involving wall-bounded flows. This model has shown good results for boundary layers subjected to adverse pressure gradients. The comparisons between the numerical simulations and the experiment results at different angles of attack are presented in Fig. 10.

The Wind tunnel (WT) validation on NACA 23012 is performed because of the availability of the experimental data. To match these data with the results of the numerical simulations the numerical setup including mesh sizing and turbulence model were the same for NACA 0012 which is easier handled during geometry variations with introducing the new morphing parameters.

## 6 Parametric study

The parametric studies using numerical simulations with ANSYS Fluent have conducted based on the results previously obtained by the X-FOIL imported to MATLAB. The morphed airfoil shapes used in the numerical simulations have selected corresponding to the maximum important aerodynamics coefficients of the airfoil such \( C_{{L_{MAX} }} ,\; \left( {C_{L} /C_{D} } \right)_{MAX} \) and maximum stall angle. The morphed shapes characteristics are compared with the basic shape of NACA 0012. The parametric study performed for each variable alone that gives a modified shape of the basic airfoil corresponding to the maximum value for this parameter.

## 7 Maximum stall angle *α*

*C*

_{D}versus angle of attack

*α*for basic and morphed airfoils. The morphed contribution at high angle of attacks gives lower coefficient of drag values. Figure 14 shows

*C*

_{L}/

*C*

_{D}ratio for basic and morphed airfoil the maximum

*C*

_{L}/

*C*

_{D}for basic airfoil is 49.8 at 10° angle of attacks while the morphed airfoil gave 40.2 at 16° angle of attack.

## 8 Maximum lift to drag ratio *C* _{L}/*C* _{D}

*C*

_{L}/

*C*

_{D}= 52.9 at angle of attack equals 10° increased by 6.22% compared to basic airfoil Fig. 20.

## 9 Maximum lift coefficient \( C_{{L_{MAX} }} \)

*C*

_{L}to its maximum value equal

*C*

_{L}= 1.34 at 18° stall angle of attack while the basic airfoil gave coefficient of lift equal

*C*

_{L}= 1.23 at lower stall angle of attack equal 16° increased by 11% from the basic airfoil the data of both airfoils shown at Fig. 24.

## 10 Controlled *C* _{L} using morphed airfoils

*C*

_{L}at different angle of attack are shown in Fig. 29.

In order to maximize the performance of the wing at all flight regimes (different angle of attacks) the airfoil shape of the wing will be changed according to its angle of attack to give the maximum certain aerodynamic characteristics such as (*C*_{L}) that is called controlled *C*_{L}.

*C*

_{L}comparison between basic and controlled

*C*

_{L}airfoils at every angle of attack. The result has indicated that controlled

*C*

_{L}morphed airfoil gave a better aerodynamic characteristics compared to basic airfoil at high angle of attacks.

## 11 Conclusions

The current study has introduces the variations in aerodynamic characteristics due to airfoil morphing including droop nose and trailing edge morphing. The morphed airfoil configuration parameterization facilitates the parametric study. The combined effect of droop nose and trailing edge morphing was studied. The results from the X-FOIL showed increase in stall angle up to 26° and lift to drag ratio up to 120. The numerical simulation study with ANSYS performed which targeted different important aerodynamic parameter improvement. From the maximum stall angle point of view the results showed increase of stall angle up to 20° and lift to drag ratio increased up to 6.22% compared to basic airfoil at separate morphed shapes. Also the separation has delayed to 75.12% of the chord length. While for the basic airfoil the separation starts at 66.4%. From the maximum Lift to drag ratio *C*_{L}/*C*_{D} point of view the morphed airfoil gave the maximum value for *C*_{L}/*C*_{D}= 52.9 at angle of attack equal 10° increased by 6.22% from the basic airfoil. Also the separation has delayed to 84.14% of the chord length. From the maximum lift coefficient point of view the lift coefficient increased also up to 1.34. Also the separation has delayed to 83.39% of the chord length. This study has showed a level of success while controlling the airfoil morphing associated with certain aerodynamic condition such maximum lift coefficient during the variety of operating angle of attack.

## Notes

### Compliance with ethical standards

### Conflict of interests

The authors declare that they have no conflict interest.

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