Week 7 - Mini project - Flow over an Airfoil

MINI PROJECT – FLOW OVER AN AIRFOIL

AIM:

To simulate an airfoil for different angle of attacks and calculate the drag and lift coefficients.

INTRODUCTION:

The airfoil simulation done here helps us to understand the lift forces that could be generated by modelling an airfoil, thereby could be used to different applications including airplanes and race cars.The airfoil which is used for simulating is the NACA 2412 airfoil. The data points are taken from airfoil tools website and designed in Converge CFD using vertex and patch method.

Then, the airfoil is simulated for different angle of attacks -  0 degree ,5 degree, 10 degree , 15 degree and calculations are done to calculate the drag and lift forces.

CASE SETUP:

The inlet velocity is calculated  by fixing the reynold’s number as 200000. From this,

`Re = (ρVD) / μ`

Where, ρ  is the density of air (1.184 kg/m³) at 26 degree C

V is the inlet velocity,

D is the characteristic length (1 metre),

μ is the dynamic viscosity of air(1.849 x 10^-5 kg/m-s).

From this, the velocity can be calculated as 3.12 m/s , which is given as the inlet condition.

The outlet condition is given as pressure with a value of 101325 Pa.

The side walls are assigned as 2d condition.

The top and bottom walls are given as symmetric condition.

Turbulence model – RNG k-ξ model.

Then for different angle of attacks, the airfoil is rotated with the help of rotate command, and then the same boundary conditions are applied, to simulate the flow over airfoil.

FLOW OVER AIRFOIL:

The flow over airfoil for different angle of attacks is shown below:

DRAG CALCULATION FOR AIRFOIL:

Drag is the opposite force which is generated when air is made to flow over an object. The simulation results can be used to obtain the drag force acting on the airfoil. To calculate the drag coefficient, we use,

Fd = 1/2 * ρ *Cd *A* V^2

Where F_d = Drag force,

 Is the density of air (1.184 kg/m³) at 26 degree C,

Cd is the drag coefficient,

A – Frontal area (For airfoil, the chord length is taken)

V – Inlet velocity (3.12 m/s),

By using this, the values of Cd for different angle of attacks were found.

LIFT CALCULATION FOR AIRFOIL:

Lift is the force which is acting perpendicular to the body and the value of the force determines whether the body is lifted upwards or pushed downwards. Airfoil shaped used for airplanes has positive lift characteristics( body is lifted upwards), and applications in automobiles and race cars utilise a negative lift characteristics (body pushed downwards) to improve grip during cornering of cars.

The simulation results can be used to obtain the lift force acting on the airfoil. To calculate the lift coefficient, we use,

Fl = 1/2 * ρ *Cl *A* V^2

Where F_l = Lift force,

 ρ is the density of air (1.184 kg/m³) at 26 degree C,

Cl is the lift coefficient,

A – Frontal area (For airfoil, the chord length is taken)

V – Inlet velocity (3.12 m/s),

By using this, the values of Cl for different angle of attacks were found.

CONCLUSION:

The drag and lift forces for the airfoil for different angle of attacks has been found. The coefficient of lift value after 10 degrees decreases along with the separation of flow, which is called as stall angle. After this angle, the value continues to decrease.

Thus the effect of angle of attack on the drag and lift values can be understood by simulating the flow over the NACA 2412 airfoil.