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

1. Incompressible, Inviscid Model: Explain why theincompressible, inviscid model for this flow should yield liftcoefficient values that match well with experiment but will yield adrag 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 distributionsfor 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 thecorresponding boundary type that you need to select.
3. Coefficient of Pressure: Run a simulation for the NACA 0012airfoil based on the initial conditions assigned to you with a meshwith 15000 elements and a mesh with 40000 elements. Plot thepressure coefficient obtained from FLUENT on the same plot as dataobtained from experiment. The experimental data is from Gregory& O’Reilly, NASA R&M 3726, Jan 1970 and plot is provided inPDF format for you to digitize in Excel. Follow the aeronauticalconvention of flipping the vertical axis so that negative Cp valuesare above and positive Cp values are below.
4. Lift and Drag Coefficient: Obtain the lift and drag coefficientsfrom the FLUENT results on the two meshes. Compare these withexperimental or expected values (present this comparison as atable). For example, the experimental values for 10 degree angle ofattack are: Cl = 1.2219; Cd = 0.0138.