NACA0012 airfoil compressibleΒΆ

NOTE: Before running this case, please read the instructions in NACA0012 airfoil incompressible to get an overall idea of the DAFoam optimization setup.

This is an aerodynamic shape optimization case for an airfoil at transonic conditions. The summary of the case is as follows:

Case: Airfoil aerodynamic optimization
Geometry: NACA0012
Objective function: Drag coefficient
Design variables: 40 FFD points moving in the y direction, one angle of attack
Constraints: Symmetry, volume, thickness, and lift constraints (total number: 81)
Mach number: 0.7
Reynolds number: 2.3 million
Mesh cells: 8.6K
Adjoint solver: rhoSimpleCDAFoam

The configuration files are available at Github. To run this case, first source the DAFoam environment (see Tutorials). Then you can go into the run folder and run:

./ 1

The optimization progress will then be written in the log.opt file.

For this case, the optimization converges in 17 steps, see the following figure. The baseline design has C_D=0.01777, C_L=0.5000, and the optimized design has C_D=0.01205, C_L=0.5000.


In this case, we need to use rhoSimpleCDAFoam, a compressible solver that uses the SIMPLEC algorithm. The case setup is similar to NACA0012 airfoil incompressible. The major difference is in the aeroOptions dictionary where we need to define different divschemes, fvrelaxfactors, and simplecontrol. These parameters are critical to ensure robust flow simulations for transonic conditions.