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Importance of wings in aircraft

An aircraft is made up of many parts that work together to propel the aircraft through the air; each part serves a specific purpose, the most important of which is to generate lift again for aircraft and control its movement according to the pilot's instructions. A significant number of moving parts are used for turning, climbing, and landing operations. The Analytical method has provided an easier way to study the aerodynamical behavior of objects under fluid motion. The Computational Fluid Dynamics (CFD) is changing the world of aviation today and this numerical simulation method allows us to make continuous evolution of aviation technology and other fields as well. Traditionally large wind tunnels were used to study the aerodynamics of objects and now it could easily do through simulation software.

Wings are the most important component of aircraft and responsible for 75% of lift generation for the flight. The wings are simply a surface fabricated along a specified 2D airfoil shape through a certain length and when this surface moves through the air certain aerodynamical forces are generated, the force which is perpendicular to the direction of motion is known as lift force. Wings are the best solution for obtaining lift and another main force is drag hence Lift-drag ratio is an important factor for the aircrafts wing. The angle produced between the air entering the wings and the chord length of an airfoils is known as an Angle of Attack (AoA)..

Airfoil. Any aerodynamic structure's cross-sectional profile is utilized to analyze its aerodynamic qualities. Symmetrical and asymmetrical airfoils are the two basic types of airfoils. When there is a need for further lift to drag (i.e., the L by D ratio must be larger), an asymmetric airfoil is utilized.

The most common airfoil used on an airplane wing is an asymmetric airfoil. It's commonly a flat-bottomed airfoil with a curved top. As a result, when air travels through the top, it moves faster than the air just at base. An area with low pressures forms on a top as the result of this. Air goes from a high pressure area to the low pressure area.

Airfoil parameters and aerodynamic forces

Figure 1: Airfoil parameters and aerodynamic forces

When liquid passes over an object's surface, it exerts force on the object. The drag force perpendicular to the lifting force is described as a force that happens right angles to the direction of the entering flow. These forces are known as aerodynamic forces when the environment is air.

NACA 0012 airfoil coordinates

Figure 2: NACA 0012 airfoil coordinates

The profile is defined by the NACA 4 wing sections:

  1. In the first number, maximum camber is expressed as a percentage of the chord.
  2. In tenths of a chord, the second digit represents the greatest camber initially from the leading edge of the airfoil.
  3. An airfoil's maximum thickness like an percentages of the chord is shown by the final two figures.
  4. For example, an NACA 2412 airfoil's has an maximum camber of 2% at 40 percent (0.4 chords) from the leading edge and a maximum chord thickness of 12%.

A NACA 0012 airfoil is symmetric, with no camber indicated by the 00. The number 15 denotes a 15 percent width to chord length ratio, meaning the airfoil is 15 percent as thickness as it is long.

Many different studies on aerodynamic force analysis can be found in the literature. The NACA 0012 airfoil was used to mimic and replicate computational fluid dynamics processes. ANSYS Fluent software was used to run analyses to determine the pressure and velocity dispersion on the wing surface. In addition, variable relative velocities were uses for derive a drag & lift coefficient. The acquired results were found to be line with the theory results. Dynamic analysis is a sort of analysis used to figure out how much stress is generated by the applied loads. Stresses owing to frequencies were determined using random vibration signals and modal analysis. The load factor and deformation were calculated using the buckling analysis. To estimate drag and lift forces, a CFD study was used. A drags coefficient were modifies superior than an actual models, according to the results of the investigation. CFD analysis and wind tunnel testing may lift and drag forces may both be calculated using both of these methods. A NACA 0012 airfoil’s was subjected to the two-dimensional low velocity flow analysis investigation in various AoAs and at 3E+ 6 Reynolds number. The outcomes of the analysis and simulation are identical. As a result, the analyses demonstrated their reliability as a substitute for experimental methods..

Airfoil and aerodynamic forces

We are using ANSYS fluent software 2021 is used for this study and we are going to create simulation for low-speed aerodynamic behavior for NACA 0012 airfoil by fluctuating the angles for attacks from 0 to 20 degree. A mesh created has a node in total 69075 which are triangular in shape and then CFD analysis was done, the turbulence models employed in the study were the K-omega turbulent system, Standard K-Epsilon Turbulence Model and Spalart Allmaras Turbulence Model When the analytical results were compared to the physical wind tunnel data, it was determined that they were now all in great agreement.

CFD simulation in fluid dynamics and aerodynamics uses mathematical models and techniques to solve issues involving fluid flows. This entails using simulation software to compute related equations determined by science, Using Navier stokes equation the simulation interface was created by applying the boundary conditions in the ANSYS simulation platform according to the Navier stokes equation. 

An unstructured mesh with a sphere of influence centered on the center of the airfoil was chosen for discretization of the finite element model. Figure 2 shows the mesh that was used in the analysis. For the analysis, a pressure-based steady-state solver with a realizable turbulence model was used.

The ANSYS analysis parameters for NACA 0012 are listed below in the table: -

Number of nodes

11506

Number of elements

11506

Solver

Pressure based steady state

Viscous model

k-epsilon

Density

1.225

Inlet velocity

1

Turbulence viscosity

10

Momentum

Second order upwind

Chord-length

1

Meshing of simulated wing model

Figure 3: Meshing of simulated wing model

Numerical computations were performed using the ANSYS FLUENT programme, which is built on the finite elements approach. Gravity should be ignored. Pressure-based and steady-state solutions were used in the calculations. The turbulence model is Standard k-e, Regular Wall Func.

Second-order functions are used to model momentum. The usage of higher order algorithms improves the results' accuracy. However, processing time is longer for high-order function solutions than for low-order function solutions. For different regions, various boundary conditions could be calculated. The density of air is 1.225 kilograms per cubic meter, and its kinematic viscosity is 1.78e-5. A 1 m/s input velocity was assumed. Simulations are performed using the above-mentioned parameters.

The darkened area illustrates how the boundary layer reduces flow. If the surface of a viscous incompressible or frictionless fluid is moved outward by the distance, where is the displacement thickness, the flow of the fluid will be decreased by the same amount. A controller that root region is functioning in a boundary layer can use the displacement thickness to minimize its effectiveness span or effective aspect ratio. This could include a rudder working under a reasonably flat hull or a roll-stabilizer fin acting near the bilge circle in the hull’s boundary layer.

Similarly, in mathematical simulations involving fluid flow with presumed inviscid flow and thus no boundary layer, the body's surface can be moved outwards by to generate a body form that is equal to that without a boundary layer. The following are some rough estimations of displacement thickness: Flow in a laminar direction:

Boundary Layer Separation 

Boundary condition 

Figure 4: Boundary condition 

Our purpose of this study is to obtain the lift and drag coefficients for the given airfoil NACA 0012 through CFD analysis and calculate the lift-drag and other parameters with changing angle of attack. Results show the static pressure measurements at 0 degree and 20-degree AoA values, respectively. The velocity results are shown at 0 degree and 20-degree Angle - of - attack values, respectively. The airfoil is also optimized using NACA 0012 as a 10 percent, 20 percent, 30 percent, and 40 percent improvement. Their CL, CD, and CL/CD findings are also provided.

Simulation process

Aerodynamic contour for pressure distribution at 0-degree Angle of attack

Figure 5: Aerodynamic contour for pressure distribution at 0-degree Angle of attack

Pressure distribution surrounding the airfoil geometry and it could be seen maximum pressure distribution on the leading edge of the airfoil. At zero-degree angle of attack and symmetric results are there in the contour.

 Aerodynamic contour of pressure distribution at 20-degree Angle of attack

Figure 6:Aerodynamic contour of pressure distribution at 20-degree Angle of attack

Pressure distribution surrounding the airfoil geometry and it could be seen maximum pressure distribution on the below leading edge of the airfoil. At 20-degree angle of attack and symmetric results are there in the contour [6].

 Velocity distribution contour at 0-degree Angle of attack

Figure 7: Velocity distribution contour at 0-degree Angle of attack

Velocity distribution along the airfoil in fluid flow at zero-degree angle of attack and maximum velocity is seen above and below of the airfoil minimum at the leading edge.

 Velocity distribution contour at 20-degree Angle of attack

Figure 8: Velocity distribution contour at 20-degree Angle of attack

Velocity distribution along the airfoil in fluid flow at 20-degree angle of attack and maximum velocity is seen above the airfoil minimum at the trailing edge. An unsymmetric contour is formed.

 Graphical representation between CL and Angle of attack

Figure 9: Graphical representation between CL and Angle of attack

To use Simpson’s rule, we will for the given graph under a closed interval we will divide the curve into even number of divisions for our case we divided it into 2 intervals.

Next find a quadratic approximation of the area enclosed under the curve.

a quadratic approximation

For our case to find Cl

yL=0

yC=3.5

yR=5

h=10

2

Graphical representation between CD and Angle of attack

Figure 10:Graphical representation between CD and Angle of attack

Graphical representation between CL/CD and Angle of attack

Figure 11: Graphical representation between CL/CD and Angle of attack

Graph plot between Cp and x/c parameter

Figure 12: Graph plot between Cp and x/c parameter

Conclusion

Airfoil are indeed the main lift producing unit of an airplane that allows it to overcome its weight. The aerodynamic forces operating on the aircraft are determined by the form of the airfoil, while drag is the force acting in the opposite direction of motion. Because these two forces are so crucial in aircraft flight, the shape of the airfoil determines the aircraft's aerodynamic performance. The aerodynamic properties of an airfoils from the NACA 0012 series, which is widely used in aircrafts, will be highlighted in this section. This study will be conducted using CFD on ANSYS 2021, as well as the graphical results and computation contour lines will be examined to determine the relations between various aerodynamic parameters.

When the findings were examined, it was discovered that at 50, 100, and 150 AoA, between 5percent and 10percent improvement was attained. At 200 AoA, improvements of ten percent and fifteen percent were attained. At all AoA values, there is essentially no increase in CD values. CL/CD to CL values experienced similar changes. Because lift force was not generated for symmetric airfoils at 00 AoA values, no sufficient modifications were observed for all 00 AoA values.

References 

[1] Ganesh Ram, R. K., Cooper, Y. N., Bhatia, V., Karthikeyan, R., & Periasamy, C. (2017). Design optimization and analysis of naca 0012 airfoil using computational fluid dynamics and genetic algorithm. In Applied Mechanics and Materials (Vol. 664, pp. 111-116). Trans Tech Publications Ltd.

[3] Kurtulus, D. F. (2015). On the unsteady behavior of the flow around NACA 0012 airfoil with steady external conditions at Re= 1000. International Journal of Micro Air Vehicles, 7(3), 301-326.

[4] Iqbal, M. Y., Shah, S. I. A., & Hassan, A. (2019, August). CFD Analysis of NACA-0012 Airfoil with Various Porous Gurney Flap Geometries. In 2019 International Conference on Applied and Engineering Mathematics (ICAEM) (pp. 231-236). IEEE.

[5] Sandeep, D., Ravitej, Y. P., Khot, S., Ravikumar, R., Abhinandan, Kumar, N., ... & Veerachari. (2021, February). CFD simulation of transonic turbulent flow past NACA 0012 aerofoil. In AIP Conference Proceedings (Vol. 2316, No. 1, p. 020013). AIP Publishing LLC.

[6] Abobaker, M., Elfaghi, A. M., & Addeep, S. (2020). Numerical Study of Wind-Tunnel Wall Effects on Lift and Drag Characteristics of NACA 0012 Airfoil. CFD Letters, 12(11), 72-82.

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[Accessed 03 March 2024].

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