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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(naca651212-il) NACA 65(1)-212 | NACA 65(1)-212 airfoil Max thickness 12% at 40% chord Max camber 1.1% at 50% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca651212-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca651212-il | 50,000 | 9 | 26.2 at α=6° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca651212-il | 50,000 | 5 | 28.4 at α=5.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca651212-il | 100,000 | 9 | 46.8 at α=4.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca651212-il | 100,000 | 5 | 42.2 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca651212-il | 200,000 | 9 | 55.9 at α=4.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca651212-il | 200,000 | 5 | 52.9 at α=4° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca651212-il | 500,000 | 9 | 76.3 at α=3.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca651212-il | 500,000 | 5 | 66.1 at α=3° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca651212-il | 1,000,000 | 9 | 87.8 at α=3.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca651212-il | 1,000,000 | 5 | 70.8 at α=2.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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