Flow simulation over a NACA 0017 airfoil for different angle of attacks in Solidworks

Aim:

To model the flow over a NACA 0017 airfoil for different angle of attacks.

Objective:

• To simulate the flow over a NACA 0017 airfoil.
• To compare the lift and drag properties for various angle of attacks.

Theory:

An airfoil (American English) or aerofoil (British English) is the cross-sectional shape of a wing, blade (of a propeller, rotor, or turbine), or sail.

An airfoil-shaped body moving through a fluid produces an aerodynamic force. The component of this force perpendicular to the direction of motion is called lift. The component parallel to the direction of motion is called drag.

The lift on an airfoil is primarily the result of its angle of attack. When oriented at a suitable angle, the airfoil deflects the oncoming air (for fixed-wing aircraft, a downward force), resulting in a force on the airfoil in the direction opposite to the deflection. This force is known as aerodynamic force and can be resolved into two components: lift and drag.

A stall is a condition in aerodynamics and aviation such that if the angle of attack increases beyond a certain point, then lift begins to decrease. The angle at which this occurs is called the critical angle of attack. This angle is dependent upon the airfoil section or profile of the wing, its planform, its aspect ratio, and other factors, but is typically in the range of 8 to 20 degrees relative to the incoming wind ("relative wind") for most subsonic airfoils. The critical angle of attack is the angle of attack on the lift coefficient versus angle-of-attack curve at which the maximum lift coefficient occurs.

Stalling is caused by flow separation which, in turn, is caused by the air flowing against a rising pressure. As angle of attack increases, the separated regions on the top of the wing increase in size as the flow separation moves forward, and this hinders the ability of the wing to create lift. Beyond the critical angle of attack, separated flow is so dominant that additional increases in angle of attack cause the lift to fall from its peak value.

Setup Explanation:

1. First, the airfoil profile is modelled by importing the coordinates of the airfoil under the curves option. Make sure the unit system is in MKS. The airfoil co-ordinates can be found over here: http://airfoiltools.com/airfoil/naca4digit?MNaca4DigitForm%5Bcamber%5D=0&MNaca4DigitForm%5Bposition%5D=0&MNaca4DigitForm%5Bthick%5D=17&MNaca4DigitForm%5BnumPoints%5D=81&MNaca4DigitForm%5BcosSpace%5D=0&MNaca4DigitForm%5BcosSpace%5D=1&MNaca4DigitForm%5BcloseTe%5D=0&yt0=Plot
2. The angle of attack can be set by rotating the sketch components. Extrude the model to an appropriate length.
3. Create separate cad files for each angle of attack and save them.

4. Then, under flow simulation option click on wizard tool and create a new setup with default options, external flow study with air as the working fluid. Set velocity in x-direction as 200 m/s (subsonic flow).
5. Edit the computational domain size such that it has 5 meters of flow before the airfoil and 15 meters of flow after the airfoil as shown. Select the 2-D simulation option under type of domain.

6. Keep the mesh setting to 3 and add a local mesh on the curved surface with a refinement level of 3 for higher accuracy.
7. Insert surface goals of force in x-direction (drag) and force in y-direction (lift) on the curved surface of the airfoil.
8. Insert a velocity cut plot parallel to the flow.
9. Insert the goal plots as well.
10. Do the above processes for angle of attack of 0, 2, 4, 6, 8 and 10 degrees.

Results:

• At 0 degrees:

• At 2 degrees:

• At 4 degrees:

• At 6 degrees:

• At 8 degrees:

• At 10 degrees:

• Variation of lift and drag:

Conclusions:

1. From 0-8 degrees the flow is streamline and the lift increases linearly.
2. After 8 degrees there is flow separation indicated by the blue region in the plot. Hence the lift decreases as well as indicated in graph.
3. The drag also steadily increases with increase in angle of attack due to increase in frontal area.

File Link:

https://drive.google.com/drive/folders/13ftVvyTOHeN0ziR4IKjOYgXedxeP-R8x?usp=sharing