Exercises

1. Alternate Meshing Strategy Using "inflation" and "sphere of influence" 

Watch the following video for a demonstration on how to create a mesh for the airfoil using "inflation" and "sphere of influence" features in the ANSYS mesher. The advantage of this meshing strategy is that it can be adapted to complex geometries. 
 

2. Airfoil CFD Validation Using a NASA-provided Mesh.  

Click here for a more detailed description. 

3. Angle of Attack Variation

Consider the low-speed airflow over the NACA 0012 airfoil at low angles of attack. The Reynolds number based on the chord is Rec = 2.88 × 10^6. This flow can reasonably be modeled as incompressible and inviscid. Explain why the incompressible, inviscid model for this flow should yield lift coefficient values that match well with experiment but will yield a drag coefficient that is always zero.

Boundary Value Problem

What is the boundary value problem (BVP) you need to solve to obtain the velocity and pressure distributions for this flow at an angle of attack of 10 degrees? Indicate governing equations, domain and boundary conditions (u = 0 at a certain boundary etc.). For each of the boundary conditions, indicate also the corresponding boundary type that you need to select in FLUENT.

Coefficient of Pressure

Run a simulation for the NACA 0012 airfoil at angles of attack at 0 degrees and 10 degrees for two cases: a mesh with 15000 elements and a mesh with 40000 elements. Plot the pressure coefficient obtained from FLUENT on the same plot as data obtained from experiment The experimental data is from Gregory & O’Reilly, NASA R&M 3726, Jan 1970 and is provided here Follow the aeronautical convention of flipping the vertical axis so that negative Cp values are above and positive Cp values are below. This can be done in MATLAB using set(gca, ’YDir’, ’reverse’);

Lift and Drag Coefficient

Obtain the lift and drag coefficients from the FLUENT results on the two meshes. Compare these with experimental or expected values (present this comparison as a table). The experimental values for 0 degree angle of attack are: Cl = 0.025; Cd = 0.0069, and the experimental values for 10 degree angle of attack are: Cl = 1.2219; Cd = 0.0138.

Conclusions

Comment on the comparison with experiment for the two angles of attack. Also,comment on the effect of mesh refinement. How does the pressure distribution over the airfoil change on increasing the angle of attack?

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