Flow over an airfoil

Title of the Project: Flow over an aerofoil

Aim: To study the flows over an aerofoil at different angle of attacks and the effects of angle of attack, the variations of lift and drag with angle of attacks.

Description and Procedure: The referred is taken form the referencce mentioned below and the aerofoil is structured anthe flow analysis is simulated with the conditions in 2-D plane.

Velocity : 600 m/s

Time dependent flow with a step time of 1 sec.

The parameters in the numerical code can be entered into equations to precisely generate the cross-section of the airfoil and calculate its properties.

Here , the NACA 0017 airfoil has a maximum camber of 0% located 0% (0 chords) from the leading edge with a maximum thickness of 17% of the chord.

Here, there is no camber(bulge or deviation from the centre line), which means that the aerofoil is completely symmetric with the chord line.

When the angle of attack is zero:

When the angle of attack is zero, there will be no induced drag on the aerofoil.

Pressure Contour:

Velocity Contour:

When the angle of attack is 2 degrees:

Pressure contour:

Velocity Contour:

When the angle of attack is 4 degrees:

Pressure contour:

Velocity Contour:

When the angle of attack is 6 degrees:

Pressure contour:

Velocity Contour:

When the angle of attack is 8 degrees:

Pressure contour:

Velocity Contour:

When the angle of attack is 10 degrees:

Pressure contour:

Velocity Contour:

The Lift forces and Drag forces from the simulations are formatted below:

Plot for Angle of attack vs Lift force is plotted below:

From the observation of values in the table above, we can see that there is lift force for zero angle of attack.

That means there is zero lift for an angle of attack less than zero.

Theoritical Procedure to calculate the lift and drag force:

Kinetic Energy (KE)=0.5 * Density * Free stream velocity;

Lift Force = Lift coefficient * KE * Area;

Area = chord length * Width of the aerofoil;

Chord length is the distance between the leading edge and the trailing edge;

Drag Force = Drag coefficient * KE * Area;

Generally the data is experimental from the tests.

From the lift, we can also calculate the circulation and the starting vortex using Kutta-Joukowski theorem

Inferences and results:

As the angle of attack is increased, The lift increases linearly, as shown in the plot. After a certain increase in the angle of attack, Lift starts decreasing non linearly. That condition is called stalling.

Also , the most important parameter is lift-to-drag ratio.

As the angle of attack increases, the lift-to-drag ratio first increases, reaches a
maximum, and then decreases. The maximum lift-to-drag ratio, (L/D)max, is an important
parameter in airfoil performance; it is a direct measure of aerodynamic efficiency. The
higher the value of (L/D)max, the more efficient is the airfoil.

Schematic of lift coefficient variation with the angle of attack.

Hence, For an aerofoil, Thin aerofoil , Lift and drag forces are measured and compared.

Results are explained.