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 | ||
(naca671215-il) NACA 67 | NACA 67 Max thickness 15% at 50% chord Max camber 1.1% at 50% chord | Remove Airfoil details Airfoil plotter |
(naca66210-il) NACA 66-210 | NACA 66-210 airfoil Max thickness 10% at 45% chord Max camber 1.1% at 50% chord | Remove Airfoil details Airfoil plotter |
(ncambre-il) ONERA NACA CAMBRE AIRFOIL | ONERA/Aerospatiale NACA-CAMBRE rotorcraft airfoil Max thickness 11.5% at 31.2% chord Max camber 1.4% at 15.9% chord | Remove Airfoil details Airfoil plotter |
Drawing Options
Polars for (naca671215-il,naca66210-il,ncambre-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca671215-il | 50,000 | 9 | 23.2 at α=7.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca671215-il | 50,000 | 5 | 21.4 at α=6.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca671215-il | 100,000 | 9 | 34.9 at α=5.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca671215-il | 100,000 | 5 | 25.8 at α=6.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca671215-il | 200,000 | 9 | 42.4 at α=5.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca671215-il | 200,000 | 5 | 33.4 at α=5.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca671215-il | 500,000 | 9 | 56.2 at α=4.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca671215-il | 500,000 | 5 | 55.5 at α=3.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca671215-il | 1,000,000 | 9 | 84.3 at α=3.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca671215-il | 1,000,000 | 5 | 71.3 at α=2.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca66210-il | 50,000 | 9 | 22.1 at α=5.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca66210-il | 50,000 | 5 | 27.6 at α=4.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca66210-il | 100,000 | 9 | 32.8 at α=4° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca66210-il | 100,000 | 5 | 36.7 at α=3.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca66210-il | 200,000 | 9 | 51.4 at α=3.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca66210-il | 200,000 | 5 | 46.7 at α=3.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca66210-il | 500,000 | 5 | 61.9 at α=2.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca66210-il | 500,000 | 0 | 67.5 at α=3° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca66210-il | 1,000,000 | 9 | 82.2 at α=2.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca66210-il | 1,000,000 | 5 | 67.1 at α=1.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ncambre-il | 50,000 | 9 | 26.5 at α=5.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ncambre-il | 50,000 | 5 | 29.6 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ncambre-il | 100,000 | 9 | 39.6 at α=5.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ncambre-il | 100,000 | 5 | 41.3 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ncambre-il | 200,000 | 9 | 52.9 at α=5.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ncambre-il | 200,000 | 5 | 54.3 at α=7.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ncambre-il | 500,000 | 9 | 73.9 at α=7° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ncambre-il | 500,000 | 5 | 73 at α=7.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ncambre-il | 1,000,000 | 9 | 90.4 at α=8° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ncambre-il | 1,000,000 | 5 | 88 at α=8.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| Reynolds number calculator | |||||||