Question

Consider the low-speed airflow over the NACA 0012 airfoil at low angles of attack. The Reynolds number based on the chord is roughly Rec = 2.88 × 10^6. This flow can reasonably be modeled as incompressible and inviscid. The initial input value for your simulations is provided on the bottom of this assignment. Your need to generate a report to give background introduction, and address the following issues: (inlet Velocity: 1.5 Attack angle: 5)

1. Incompressible, Inviscid Model: 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.
2. 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 any angle of attack? 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.
3. Coefficient of Pressure: Run a simulation for the NACA 0012 airfoil based on the initial conditions assigned to you with 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 plot is provided in PDF format for you to digitize in Excel. Follow the aeronautical convention of flipping the vertical axis so that negative Cp values are above and positive Cp values are below.
4. 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). For example, the experimental values for 10 degree angle of attack are: Cl = 1.2219; Cd = 0.0138.

Answer 1