Compressible Flow: NACA 0012 Airfoil

This project simulates the flow over a NACA 0012 airfoil using ANSYS Fluent, with compressible flow as the central modelling theme. At the freestream conditions studied here, the air can no longer be treated as incompressible — density varies appreciably with pressure and temperature across the flow field — so the simulation is built around a compressible-flow formulation, making it a clear illustration of how that class of flow model is set up and solved.

The airfoil is the cross-sectional shape of a lifting surface such as an aircraft wing, a wind-turbine blade or a helicopter rotor. The aerodynamic behaviour of a given design depends strongly on its profile, which is why different airfoils are selected for different applications. The geometry is defined by familiar parameters: the chord line, the leading and trailing edges, and the angle of attack — the angle between the chord and the oncoming flow direction. In this case the angle of attack is 5°, so the incoming velocity is resolved into a horizontal component of cos5° ≈ 0.996 and a vertical component of sin5° ≈ 0.087. The objective is to examine the airflow behaviour and the pressure distribution around the airfoil and to study the resulting lift and drag forces.

The geometry is created in Design Modeler and meshed in ANSYS Meshing with a structured grid of 35,000 cells.

Because the flow is compressible, a density-based solver is used — the appropriate choice when density variations are coupled tightly to the pressure and energy fields, as they are in high-speed aerodynamics. For compressible flow, the Mach number must be specified in the boundary conditions; it is the ratio of the flow speed to the local speed of sound (for reference, the speed of sound in air at 25 °C is about 343 m/s). Airfoil simulations of this kind require a far-field boundary condition with the Mach number prescribed for the surrounding flow, set here to 0.6 — firmly in the subsonic-but-compressible regime where compressibility effects are significant and cannot be neglected.

The solution produces two-dimensional contours of pressure, velocity, temperature, density and Mach number, together with streamlines around the profile. The results show the highest pressure at the leading edge, where the flow stagnates on direct contact with the airfoil, and the strongest pressure drop along the upper surface. This pressure difference between the upper and lower surfaces is what generates lift. The velocity field mirrors the pressure field exactly, as expected: regions of highest pressure coincide with the lowest velocity, and regions of lowest pressure with the highest velocity — the classic inverse relationship that underlies airfoil aerodynamics, here captured within a fully compressible treatment that also resolves the accompanying temperature and density variations.