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Efficiency of a wing is sufficient at stall angle of attack and breakdown by effecting turbulent flow intensity on upper wing surface which is destroyed the airflow pattern and generated adverse pressure, backflow airstream and forward transition separation point movement closely to the leading edge of the wing. To enhancement a wing efficiency factor at high angle of attacks in this research is using partially sucking adverse pressure to delay the separation point to control the backflow rate and expelling it through the wing tips ends using tangential suction slot channel without using any mechanical means depending on differential static pressure between adverse pressure and the static pressure generated on both win tips ends. Numerical CFD analysis has been managed the case study of the effect suction slot channel located on the upper wing surface at (70%) of the wing chord from the leading edge as a proposed modified wing model of the Cessna 172 Skyhawk half wing at angle of attacks (10°, 12°, 14°, 16°, 18°, 20°, 22°, 24°). The analysis study was conducted at stall speed (90 km/hr.) without flaps where it is equal (25 m/sec). The results of the analytical shows at the range of angle of attacks were chosen between (10° to 24°), lift coefficient CL increased (3.757%), drag coefficient CD is decreased (0.530%) and the wing efficiency lift to drag ratio Δ(CL/CD) is increased to (5.712%) while Δ(CD/CL) is decreased to (5.231%) which these parameters reflected to enhancement the wing to reduce required power at the minimum speed, increase angle of climb and decrease angle of descent reduce landing distance with high angle of attacks.
aerodynamic, tangential suction slot channel, flow control, computational fluid dynamics (CFD), aircraft half wing Cessna 172 Skyhawk
The wing efficiency is one of the most importance factor of an airplane’s aerodynamic efficiency which is measured by (lifttodrag) ratio, generating lift force at different angle of attacks is depending on the form of the airflow stream pattern passing over the surface of the wing where is closed to laminar flow at low angle of attacks on upper surface of the wing while at the high of attack the laminar pattern form is become unstable and transfer into turbulent flows which is produce higher skinfriction drag and adverse pressure gradient acting with viscous forces to reverse flow which results separated flow on the upper surface of the wing [1]. Skinfriction drag and adverse pressure gradient are affected on wing efficiency [2].
Improving wing efficiency is one of the important factors to enhancement aircraft performance to increase excessive power between available and required engine power to maintain the flight condition at critical low speed [3]. The proposed improvement to the primary sweptback wing is attained an enhancement in the maneuverability of aircraft flight and stability controls, particularly during landing, approach, descent, and take off at low speeds with the lack runway by dramatically increasing lift created at the wingtip where the efficient attack angle and the ratio of lifttodrag are enhanced as a result of smoothing the airflow across the upper wing near the tip [4]. The properties of aerodynamics and characteristics of the boundary layer are computed using the final properties data acquired from characteristics of the twodimensional airfoil at an efficient attack angle and threedimensional swirl ring forces [5].
There are many techniques to control flow stream of turbulent boundary layer on the upper surface of an airfoil in two and three dimensions at high angle of attacks using blowing upper surface of the wing to energized the boundary layer and reduce adverse pressure using part of compressed air from the aircraft engine or absorbing part of the adverse pressure from the upper surface using absorbing equipment vis tangential suction slots. The main idea of this research is to energized and control flow stream of turbulent boundary layer to increase the kinetic energy of the airflow on the upper surface by absorbing part of the adverse pressure gradient which is generated at high angle of attacks of the wing without using absorbing equipment carried out on aircraft Cessna 172 Skyhawk at high angle of attacks between (10° to 24°) as a technique to develop flow stream control. Numerical analysis by using (CFD) is achieved improvement aircraft efficiency and performance by using tangential suction slot channel on the upper surface of the wing located at 70% of the wing chord, (2.6) meters length and (3) cm slot width from the wingtip depending on the deferential static pressure magnitudes between the upper wing surface caused by adverse pressure and static pressure occurred by flow vortices on the side surface of the wingtip.
There are many investigations and techniques have been achieved to develop flow control or boundary layer control from 1904 when it employed by Prandtl [6], to boost the force of lift or reduce the force of drag as recorded in 1996 [7]. Furthermore, the combination of jet position and attack angles improved the coefficient of lift better than other suction conditions as presented in the study of Huang et al. [8]. The injection decreases the skin friction as indicated in the study of Shojaeifard et al. [9]. Meanwhile, Beliganur and Raymond [10] optimized flow control for enhancing ratio of lift to drag for a NACA 0012 airfoil, while for a NACA 0015, Jasim [11] presented a numerical simulation study of unsteady incompressible laminar flow over its airfoil with and without a continuous wallnormal suction and blowing as an active boundary layer control technique. Moreover, to improve the efficiency of highlift, a twodimensional airfoil's aerodynamic analysis to enhance the circulation control system performance as absorbed was performed by Jensch et al. [12]. In controlling the flow, the case of study is the flow field over an airfoil with four holes which are penetrated the airfoil from the upper surface to the lower surface and through which a part from the primary flow is sucked from the upper surface, and then blowing it through same holes from the lower surface. Genç et al. [13] CFD techniques were used to examine the impact of suction, blowing, and synthetic jets on the aerodynamic properties of airfoils, as well as to optimize the geometry of the suction slots and blowing on a NACA 0012 airfoil. Piperas [14] tested the control of flow separation on a NACA 4415 airfoil using various suction configurations and improved the value of maximum lift coefficient by 20%. Ravindran [15] examined the performance of tangential unsteady blowing and suction in control of flow separation on the TAU0015 airfoil. As the attack angle augments, the flow's upper surface separation point travels from the trailing edge to the leading edge [16].
Tangential suction slot on the upper surface of the wing is one of the techniques to control influence the turbulent boundary layer by partially reduction impact the adverse pressure and delay and transition forward separation point to the wing leading edge. In this approach, the effect of the tangential suction slot channel located at (70%) of the wing chord from the leading edge on flow separation from the upper wing surface at a high angle of attack is investigated. The aim of this approach is to increase the aircraft Cessna 172 wing efficiency (CL/CD) at high angle of attacks and reduce stall velocity of the aircraft. Numerical analysis using (CFD) was been accomplished through the same geometry of the main half wing of the aircraft and the proposed modification main half wing by tangential suction slot channel tested at the same flight conditions and the results are compared.
The Cessna 172 Skyhawk (Figure 1) is an American fourseat, singleengine, high wing, fixedwing aircraft made by the Cessna Aircraft Company [17], and Specifications shown in Table 1 [18].
Figure 1. Cessna 172 3view drawing
Table 1. General Aircraft Cessna 172 Skyhawk specifications
General Aircraft Cessna 172 Skyhawk specifications 

Wing Span 
11 m 
Horizontal tail span 
3.4 m 
Wing Area 
16.2 m^{2} 
Horizontal tail area 
3.74 m^{2} 
Aspect Ratio 
7.32 
H.T root chord 
1.4 m 
Taper Ratio 
0.672 
H.T tip chord 
0.8 m 
Wing Root Chord 
1.625 m 
Vertical tail area 
1.81 m^{2} 
Wing tip chord 
1.092 m 
Fuselage length 
8.28 m 
Sweep Angle 
0 o 
Power plant 
1x Lycoming IO360L2A, 4 cylinder, max hp (160), (120 kw) 
Main Aerodynamic chord 
1.47m 
Stall speed with full flaps down 
87 km/hr 
Cross section airfoil 
NACA 2412 
Stall speed without flaps 
90 km/hr 
Zero lift angle 
2.07o 
Airfoil lift slope curve 
0.105 deg^{1} 
Wing load 
68.6 kg/m^{2} 


4.1 Main half wing and proposed modified half wing of an aircraft (Cessna 172)
The main half wing of an aircraft was designed by Solid work (2015) software according to the design data and dimensions of the aircraft as shown in Figure 2. The proposed idea is to create a tangential suction slot (2.6) m length and (30) mm opening width on the upper wing surface at (70%) of the wing chord from the leading edge closed to the wingtip connected with the main suction header tube length (2.6) m and long diameter (120) mm and short diameter (80) mm is built inside the wing as shown in Figures 3 and 4, to absorb and reduce partially adverse static pressure occur in the turbulent flow over the wing at high angle of attacks and delay separation point of the reverse flow to the leading edge of the wing at stall speed without flaps. Depending on the deferential static pressure magnitudes between the upper wing surface caused by adverse pressure and static pressure occurred by flow vortices on the side surface of the wingtip without using absorption equipment.
Figure 2. Half wing geometry Cessna 172 Skyhawk
Figure 3. Tangential suction slot
Figure 4. Proposed suction tube dimensions
At high angle of attack, adverse static pressure gradient increased on the upper wing surface, separation point moves to the leading edge and breaks up airflow pattern velocity shape away from the upper surface. This is the natural phenomenon of the airflow viscous responsible increased wing drag force as affected by the skin friction during increasing angle of attack. In this approach consider the tangential suction slot one of the techniques to control backflow airstream flow by partially reduction of adverse static pressure intensity and expelling it through the wing tips ends without using any suction source depending on different static pressure magnitude on these two levels. Computer Fluid Dynamic program (CFD) has been used to investigate and study the case study the effect of tangential suction slot channel to enhance aerodynamic performance of the airplane wing at high angle of attacks with low speed.
NavierStokes for unsteady, incompressible, threedimensional viscous flow is describing the motion of viscous fluid substances where equations are three components of the momentum Equation (1), (2) and (3). However, to close the problem they also contain the continuity equation (4) according to Anderson Jr. [19]. Three components of the momentum equations are:
$\frac{\partial u}{\partial t}+\rho u \frac{\partial u}{\partial x}+\rho v \frac{\partial u}{\partial y}+\rho \omega \frac{\partial u}{\partial z}=\frac{\partial p}{\partial x}+\frac{\partial}{\partial x}(\lambda \nabla \cdot \vec{V}+$
$\left.2 \mu \frac{\partial u}{\partial x}\right)+\frac{\partial}{\partial y}\left[\mu\left(\frac{\partial v}{\partial x}+\frac{\partial u}{\partial y}\right)\right]+\frac{\partial}{\partial z}\left[\mu\left(\frac{\partial v}{\partial z}+\frac{\partial u}{\partial x}\right)\right]$ (1)
$\rho \frac{\partial v}{\partial t}+\rho u \frac{\partial v}{\partial x}+\rho v \frac{\partial v}{\partial y}+\rho \omega \frac{\partial v}{\partial z}=\frac{\partial p}{\partial y}+\frac{\partial}{\partial x}\left[\mu\left(\frac{\partial v}{\partial x}+\right.\right.$
$\left.\left.\frac{\partial u}{\partial y}\right)\right]+\frac{\partial}{\partial y}\left(\lambda \nabla \cdot \vec{V}+2 \mu \frac{\partial v}{\partial y}\right)+\frac{\partial}{\partial z}\left[\mu\left(\frac{\partial w}{\partial y}+\frac{\partial v}{\partial z}\right)\right]$ (2)
$\rho \frac{\partial w}{\partial t}+\rho u \frac{\partial w}{\partial x}+\rho v \frac{\partial w}{\partial y}+\rho \omega \frac{\partial w}{\partial z}=\frac{\partial p}{\partial z}+\frac{\partial}{\partial x}\left[\mu\left(\frac{\partial u}{\partial z}+\right.\right.$
$\left.\left.\frac{\partial w}{\partial x}\right)\right]+\frac{\partial}{\partial y}\left[\mu\left(\frac{\partial w}{\partial y}+\frac{\partial v}{\partial z}\right)\right]+\frac{\partial}{\partial z}\left(\lambda \nabla \cdot \vec{V}+2 \mu \frac{\partial w}{\partial z}\right)$ (3)
The continuity equation is given by:
$\frac{\partial \rho}{\partial t}+\frac{\partial(\rho u)}{\partial x}+\frac{\partial(\rho v)}{\partial y}+\frac{\partial(\rho w)}{\partial z}=0$ (4)
Lift and drag coefficient can be obtained by [6, 9]:
$C_{L}=\frac{1}{C} \int_{0}^{C}\left(C_{p, 1}C_{p, u}\right) d x$ (5)
$C_{D}=\frac{1}{C} \int_{0}^{C}\left(C_{p, 1} \frac{d y}{d x}C_{p, u} \frac{d y}{d x}\right) d x$ (6)
$C_{P}=\frac{PP_{\alpha}}{\frac{1}{2} \rho v^{2}}$ (7)
$R e=\frac{V_{\infty} C \rho_{\infty}}{\mu}$ (8)
Computer Fluid Dynamic (CFD) was performed to numerical analysis in this approach to calculate coefficients of lift and drag (C_{L} and C_{D}) respectively, (C_{L}/C_{D}) and (C_{D}/C_{L}) at different high angle of attacks between (10°24°). Spallart Allmaras' turbulence model was selected; it is a one equation model developed exclusively for applications of aerospace; It solves a kinematic eddy viscosity transport equation [20]. To discretize the equations of the governing, a secondorder upwind technique was used, and the SIMPLE method was employed to solve them.
7.1 Boundary condition
Numerical simulation of both half wings was conducted under the same boundary condition. Airflow boundary fluent enclosure has been created by dimension (X=18 m, Y=12 m, and Z=7 m) as shown in Figure 5 to avoid the boundary layers effect on stream flow according to study of Attia et al. [21].
Figure 5. Airflow boundary enclosure
where, the boundary condition of the numerical analysis was performed by airstream:
7.2 Mesh generation and grid independence test
The nodes number and elements number affect the results accurateness and the required time for the simulation. The larger the number of cells, needs more computing time and the longer convergence time required. The main basic half wing model of the Cessna 172 aircraft was meshed by ANSYS workbench software through using tetrahedrons method, inflation, Curvature and Patch independent. On a basic wing model with (0°) angle attack, a grid independent test (GIT) for the mesh was performed to select the appropriate number of elements and nodes in order to obtain highly precise outcomes and reducing simulation time and which would then be applied on the models of the basic wing and a modified wing with different attack angles. Six various meshing cases were opted to evaluate the Grid independent test as shown in Figure 6 [22]. The GIT for the six cases was attained by comparing the results of drag coefficient (CD) on the basic wing model according to the nodes number and elements number as shown in Table 2. The results of six cases are plotted as shown in Figure 7, was observed match results of the drag coefficient (CD) according to the number of elements for the cases (3, 4, 5, and 6). Case (3) was selected to provide precise findings while decreasing simulation period depending on the significant effect on the Central Processing Unit (CPU) of computational time which is Figure 8 as mentioned in Ref. [23].
Table 2. The elements and nodes number applied in the verification of the GIT
Case Number 
Nodes 
Elements 
CD 
Case (1) 
228708 
823143 
0.00288 
Case (2) 
414690 
1481840 
0.01221 
Case (3) 
516077 
1949924 
0.01193 
Case (4) 
618075 
2266161 
0.01210 
Case (5) 
701033 
2655588 
0.01191 
Case (6) 
802274 
3129037 
0.01190 
Figure 6. Meshing cases for conducting the GIT
Figure 7. Grid independence test results for drag coefficient main half wing
Figure 8. Case (3) Meshing detail of the model
Numerical Analysis was been performed on the basic half wing and the proposed modification of the aircraft. The results of the analytical simulation are shown in Table 3 which are represent drag coefficient (CD), lift coefficient (CL), wing efficiency as a ratio of lift to drag coefficients, these parameters are affected by increasing angle of attacks. The effect of the suction slot on the lift and drag coefficients at high angle of attacks between (10° and 24°) as a selected angle of attacks at the stall velocity without using flaps is (25 m/s). The suction slot affected to increase aerodynamic efficiency of the wing by partially sucking adverse backflow pressure and delayed forward movement of separation point than the original main wing. Figures 912 are graphs of represents relationship between the CL, CD as a lift and drag coefficients respectively, and CL/CD and CD/CL are ratio lifttodrag coefficients as a wing efficiency at a selected angle of attacks respectively, Figure 13 is plot of relationship between CL and CD which is represent the drag polar of the wing.
Figure 9. Lift Coefficient CL Vis AoA
Figure 10. Drag Coefficient CD Vis AoA
Figure 11. Wing Efficiency CL/CD Vis AoA
Figure 12. Wing Efficiency CD/CL Vis AoA
Figure 13. Drag Polar CL Vis CD
As demonstrated in Figures 913 and Tables 36, when the angle of attack is increased, lift Coefficient (CL), drag coefficient (CD), and wing efficiency (CL/CD) increase, while the (CD/CL) reduces. Moreover, the percentage rates ratio of change values of the CL between (10 to 24) degrees of angle of attacks are (3.757%) and the percentage rates ratio of change values of the CD between (10 to 24) degrees of angle of attacks is (0.530%).
Table 3. Drag coefficient, lift coefficient and the ratio of lift to drag for the both wing models
AoA (Degree) 
Basic Half wing 
Modified Half wing with slot 

CL 
CD 
CL /CD 
CL 
CD 
CL / CD 

10 
0.97795 
0.05439 
17.98032727 
1.0092 
0.05413 
17.98032727 
12 
1.14058 
0.06964 
16.3782309 
1.16163 
0.06933 
16.52056779 
14 
1.26501 
0.08596 
14.71626338 
1.29312 
0.08562 
14.81970478 
16 
1.37827 
0.10396 
13.25769527 
1.40112 
0.10296 
13.38646076 
18 
1.44739 
0.12387 
11.68475014 
1.49508 
0.12275 
11.68475014 
20 
1.52708 
0.15196 
10.04922348 
1.56536 
0.14693 
10.04922348 
22 
1.54828 
0.18662 
8.296431251 
1.60912 
0.18054 
8.296431251 
24 
1.48611 
0.234 
6.090614754 
1.53659 
0.21527 
6.350897436 
Table 4. Represent a change of values CL, CD, CL/CD and CD/CL between the two models indicated as ΔCL. ΔCD, Δ(CL/CD) and Δ(CD/CL) depending on a selected angle of attacks
AoA (Degree) 
ΔCL 
ΔCD 
Δ(CL/CD) 
Δ(CD/CL) 
10 
1.031955 
0.99522 
1.036911 
0.9644 
12 
1.018456 
0.995549 
1.02301 
0.97751 
14 
1.022221 
0.996045 
1.02628 
0.97439 
16 
1.016579 
0.99471 
1.02198 
0.97849 
18 
1.032949 
0.990958 
1.04237 
0.95935 
20 
1.025067 
0.966899 
1.06016 
0.94325 
22 
1.039295 
0.96742 
1.07429 
0.93084 
24 
1.033968 
0.882254 
1.17196 
0.85327 
Table 5. Represents the percentage change in parameter values due the effect tangential suction slot for a selected angle of attacks
AoA (Degree) 
ΔCL% 
ΔCD% 
Δ(CL/CD)% 
Δ(CD/CL)% 
10 
3.125 
0.026 
3.691 
3.556 
12 
2.105 
0.031 
2.301 
2.249 
14 
2.811 
0.034 
2.628 
2.561 
16 
2.285 
0.055 
2.198 
2.151 
18 
4.769 
0.112 
4.237 
4.065 
20 
3.828 
0.503 
6.016 
5.675 
22 
6.084 
0.608 
7.429 
6.916 
24 
5.048 
2.873 
17.196 
14.673 
Table 6. Represent the percentage ratio of change parameter values
ΔCL % 
ΔCD % 
Δ(CL/CD)% 
Δ(CD/CL)% 
3.757 
0.530 
5.712 
5.231 
Figure 14. Absorption effect of the tangential sucking slot on adverse pressure
AoA 
Main half wing Pressure Contour 
Modified half wing Pressure Contour 
Scale 
10 

12 

14 

16 

18 

20 

22 

24 
Figure 15. Pressure Contour
The difference in static pressure arising between the upper surface of the wing at high angle of attacks because of the adverse pressure and the tip of the wing led to absorption partially backflow of the air stream from the upper surface through the suction header tube by the tangential suction slot to discharge it through the ends of the wing, as a resultant that reduction partially adverse pressure and restricted forward movement of the separation to the leading edge as shown in Figure 15. The suction flow is shown in Figure 14 due to the crosssection plan created by the CFD in section the suction header tube for a chosen angle of attack between (10° to 24°).
In this research, the case study was implemented on a half wing geometry of aircraft Cessna 172. To improve wing efficiency at limited flight condition with critical stall speed (90 km/hr.) and stall angle of attack (24°) without using lift devices (flaps) according to the aircraft data, to maintain a steady level flight. Numerical analysis was performed using (CFD) software to analyze influence tangential suction slot on the upper surface at 70% of the chord measuring from the leading edge to energized the kinetic energy to control the airflow on the upper surface of the wing in a turbulent boundary layer by absorbing part of the adverse pressure generated at high angle of attacks physically without absorbing machines or equipment depending on the deferential static pressure magnitudes between the upper wing surface caused by adverse pressure and static pressure occurred by flow vortices on the side surface of the wingtip The effect of this technique on the lift coefficient CL and drag coefficient CD magnitudes as a main parameter to determine the wing efficiency as a liftto drag ratio and wing drag polar as a relationship between lift coefficient to the drag coefficient to improve the wing performance at high angle of attacks between (10° to 24°) degree where the angle of attack (24°) is a stall angle of attack. The results show the percentage rates ratio of change values in lift coefficient CL is increased to (3.75%) and drag coefficient CD is reduced to (0.530%) at a chosen angle of attacks between (10° to 24°) degrees, the wing efficiency ratio related to the Δ(CL/CD) is increased to (5.712%). Stall speed is reduced by 2.22% to be (88 km/hr.). The results of this research case study are improved aircraft Cessna 172 performance at critical angle of attack with stall velocity to maintain the aircraft at the straight steady flight and maneuvers with high angle of attacks and reduce landing distance.
The authors would like to thank the University of Baghdad, Iraq, for their assistance and support.
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