Flow over an airfoil

Flow Over an Airfoil

Aim:

To model the flow over NACA0017 airfoil and compare the lift and drag force for the various angle of attacks (2,4,6,8,10).

Objectives:

1. Load and model the NACA 0017 airfoil from the coordinate data text file.
2. Analyse the simulation of airfoil for different angle of attacks rom 0 to 10.
3. Compare the lift and drag of different cases.

Introduction:

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 tenths of the chord.
3. Last two digits describing maximum thickness of the airfoil as percent of the chord.^[3]

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.

Angle of attack is the angle between the line of chord of the airfoil and the relative airflow.

Profile geometry – 1: Zero-lift line; 2: Leading edge; 3: Nose circle; 4: Max. thickness; 5: Camber; 6: Upper surface; 7: Trailing edge; 8: Camber mean-line; 9: Lower surface

Geometry:

Computational Domain:

Input Parameters:

Procedure:

1. Input the airfoil curve from NACA0017.txt file(Insert > Curves >Curves through xyz points)
2. Select the spline and convert the entities and construct the line to represent as the enclosed volume.
3. Set the angle of attack by using the rotate entities.
4. Extrude the curve after converting the curve into a sketch.
5. Enable the flow simulation and select the wizard and input the respective details as per the above given table.
6. Set the computational domain to 2D and set the domain in such a way that the simulation is not affected.
7. Insert the Global goals and surface goals.
8. Edit the global mesh and define the mesh into more finer mesh.
9. Insert the local mesh on the surface of the airfoil for more accurate results.
10. Run the simulation and see the Pressure plot.
11. Plot a cut-plot for pressure and velocity contour.
12. Insert goals from result to get the drag and lift force results.

Results and Discussion:

1. At 2 degree angle of attack

Velcoity Profile:

Pressure profile:

Drag and Lift forces:

2. At 4 degree angle of attack

Velocity profile

Pressure profile

Drag and lift forces

3. At 6 degree angle of attack

Velocity profile:

Pressure profile:

Drag and lift forces:

4. At 8 degree angle of attack

Velocity profile

Pressure profile

Drag and lift forces:

5. At 10 degree angle of attack

Velocity profile:

Pressure profile:

Drag and lift forces:

6. At 0 degree angle of attack

Velocity profile

Pressure profile:

Drag and lift forces:

Conclusion:

1. As the angle of attack increases , the lift forces on the surfaces are also increased.
2. Lift force is nearly equal to zero in case of zero degree angle of attack when compared to other cases.
3. Giving the local mesh for the airfoil surface is needed to get the better results.
4. Symmetry in velocity and pressure profiles is seen in the zero degree angle of attack.

References:

1. https://en.wikipedia.org/wiki/NACA_airfoil
2. http://airfoiltools.com/search/airfoils