Flow simulation on a NACA airfoil

Objective: To demonstrate the flow simulation over a NACA airfoil at different angle of attacks

Description: The NACA airfoils are airfoil shapes for aircraft wings developed by the National Advisory Committee for Aeronautics (NACA). The shape of the NACA airfoils is described using a series of digits following the word "NACA". The parameters in the numerical code can be entered into equations to precisely generate the cross-section of the airfoil and calculate its properties.

The NACA four-digit wing sections define the profile by:

1. First digit describing maximum camber as percentage of the chord
2. Second digit describing the distance of maximum camber from the airfoil leading edge in tens of percents of the chord.
3. Last two digits describing maximum thickness of the airfoil as percent of the chord.

For example, the NACA 2412 airfoil has a maximum camber of 2% located 40% (0.4 chords) from the leading edge with a maximum thickness of 12% of the chord.

The NACA 0015 airfoil is symmetrical, the 00 indicating that it has no camber. The 15 indicates that the airfoil has a 15% thickness to chord length ratio: it is 15% as thick as it is long.

Procedure:

• Download the sample airfoil data
• Convert the spline into a sketch using the convert entities button and then join the two open ends of the spline at the tail end of the airfoil using a line
• Then we rotate the airfoil according to the different angle of attacks (0,2,4,6,8,10) using the move entities button
• Then we extrude the airfoil (in this case i have extruded to 0.2 metres)
• Build up the basic mesh and build a local mesh at the surface of the airfoil as the basic mesh is not a body fitted mesh
• Then using the flow wizard set the flow as an external flow and a transient analysis (Time dependent), fluid used is air having a density of 1.225 kg/m^3 and velocity of air in x-direction was given as 600 m/s
• Then run the simulation
• Then create a cut plot to get the velocity profile
• Then using the goal from the insert data tree get the results for global goal and surface goal i.e, lift force and drag force

Results obtained:

• At 0 degree angle of attack

Velocity profile:

Pressure profile:

Drag and lift force results:

Drag force (X): 315.903 N

Lift Force (Y): 0.645 N

• At 2 degree angle of attack

Velocity profile:

Pressure profile:

Drag and lift force results:

Drag force (X): 332.729 N

Lift Force (Y): 834.451 N

• At 4 degree angle of attack

Velocity profile:

Pressure profile:

Drag and lift force results:

Drag force (X): 414.052 N

Lift force (Y): 1643.361 N

• At 6 degree angle of attack

Velocity profile:

Pressure profile:

Drag and lift force result:

Drag force (X): 554.286 N

Lift force (Y): 2437.882 N

• At 8 degree angle of attack

Velocity profile:

Pressure profile:

Drag and lift force result:

Drag force (X): 756.038 N

Lift force (Y): 3242.122 N

• At 10 degree angle of attack

Velocity profile:

Pressure profile:

Drag and lift force result:

Drag force (X): 995.437 N

Lift force (Y): 4039.259 N

Explanation:

• At 0 degree angle of attack the pressure profile seems to be symmetric along the chord line of the airfoil
• It is seen that as the angle of attack increases the lift force also increases
• In the pressure profile of all the above cases no reflected wave from the upper and lower boundaries are seen due to increase in dimension of the computational domain from the base line dimension
• There is negligible lift force in case of zero degree angle of attack
• Mesh refinement is done by creating local mesh on the top surface of airfoil in order to get a optimal solution of the forces in X-direction and Y-direction
• Drag force tend to increase with increase in angle of attack
• Decrease in maximum pressure is observed after 2 degree of angle of attack whereas maximum pressure increases from 0 degree angle of attack to 2 degree angle of attack

References:

https://en.wikipedia.org/wiki/NACA_airfoil