Airfoil Comparison
Select airfoils from the airfoil database or add your own airfoils and compare the airfoil shape and lift/drag polars.
| Open full size plan in new window | Open paginated plan in new window | |
| Download PDF file | SVG image as text file | |
| Clear all | ||
(n0012-il) NACA 0012 AIRFOILS | NACA 0012 airfoil Max thickness 12% at 30% chord Max camber 0% at 0% chord | Remove Airfoil details Airfoil plotter |
Drawing Options
Polars for (n0012-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| n0012-il | 50,000 | 9 | 25.7 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n0012-il | 50,000 | 5 | 26.5 at α=5.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n0012-il | 100,000 | 9 | 36.7 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n0012-il | 100,000 | 5 | 36.1 at α=5.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n0012-il | 200,000 | 9 | 47.4 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n0012-il | 200,000 | 5 | 45.9 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n0012-il | 500,000 | 9 | 61.7 at α=6.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n0012-il | 500,000 | 5 | 61.7 at α=7.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n0012-il | 1,000,000 | 9 | 75.6 at α=7.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n0012-il | 1,000,000 | 5 | 75.4 at α=8.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| Reynolds number calculator | |||||||