Flow over an airfoil

Objective- Model the flow over a NACA0017 airfoil in Solidworks for the angle of attacks 0,2,4,6,8,10 compare the drag forces.

Theory- The NACA airfoils are airfoil shapes for aircraft wings. 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.

Steps-

• Download the sample airfoil Data.
• Click on insert curve then click curve through XYZ points then browse the downloaded NACA file.
• Select front plane then go to sketch then click on Convert Enities which basically convert it into sketch.
• Then using line command to close the loop of air foil.
• Then go to features and extruded it to 0.2 m to make a 3D model.

CAD model-

Flow Simulation-

• Click on flow simulation then click on wizard.
• Click on external flow and time dependent simulation.
• Air should to be consider as working fluid and adiabatic wall is the default wall thermal condition.
• Initial thermodynamics conditions are- Pressure = 101325 Pa and Temperature = 293.2 K.
• And give the velocity along x direction = 500 m/s.

Computational Domain-

• Right click on it and click on edit definition then select 2D simulation in XY plane.
• Again click on edit definition then drag the domain size-
• 5 times the chord length (= 1m) at inflow and 14 times the chord lenght at outflow. Also drag the upper and lower sides of computational domain upto some extent.

Calculation of analysis time-

Since the total length of computational domain is the sum of length at inflow + chord length + length at outflow.

Hence, total length of computational domain = 5+1+14 = 20 m.

Since, we gave the velocity = 500 m/s.

Therefore, total analysis time = Distance/ Speed

total analysis time = 20/500 = 0.4 sec.

Hence, we give total analysis time = 0.4 sec and output time stop = 0.04 sec in flow simulation.

Goals-

• Insert global goals and select force and frictional force along X-axis and Y-axis each.
• Insert surface goals and select upper surface of airfoil then select force along X-axis and Y-axis.

Mesh-

• Insert local mesh and select upper surface of airfoil and refining cells = 3 which means cells are smaller near the wall.

Move Enitities-

initially angle of attack is 0 which means airfoil is not tilted. But for other angle of attack, we can do following steps-

• Hide the computational domain and do the same steps for each angle of attack.
• Go to sketch click on move entities and select rotate entities.
• Select spline (airfoil model) and line (joining between ends of airfoil model).
• Select origin as the rotation point.
• Then make angle of attack as per per question.

• Make a cut plot in XY plane and showing mesh and select velocity as contours.
• Run the simulation.

Results-

For the angle of attacks 0 -

For the angle of attacks 2-

For the angle of attacks 4-

For the angle of attacks 6-

For the angle of attacks 8-

For the angle of attacks 10-

Conclusion-

For compare the drag force, it is directly proportional to the frictional force along X-axis

Hence the order of drag force is:

For the angle of attacks 6 > For the angle of attacks 2 > For the angle of attacks 8 > For the angle of attacks 0 > For the angle of attacks 4 > For the angle of attacks 10