Airfoil Comparison
Select airfoils from the airfoil database or add your own airfoils and compare the airfoil shape and lift/drag polars.
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(naca6412-il) NACA 6412 | NACA 6412 airfoil Max thickness 12% at 30.1% chord Max camber 6% at 39.6% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca6412-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca6412-il | 50,000 | 9 | 9.8 at α=3.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca6412-il | 50,000 | 5 | 29.6 at α=4.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca6412-il | 100,000 | 9 | 53.1 at α=10.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca6412-il | 100,000 | 5 | 58.7 at α=6.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca6412-il | 200,000 | 9 | 80 at α=8.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca6412-il | 200,000 | 5 | 81.2 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca6412-il | 500,000 | 9 | 114.2 at α=6.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca6412-il | 500,000 | 5 | 111.3 at α=6.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca6412-il | 1,000,000 | 9 | 142.7 at α=5.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca6412-il | 1,000,000 | 5 | 136.6 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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