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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(naca634221-il) NACA 63(4)-221 | NACA 63(4)-221 airfoil Max thickness 21% at 34.9% chord Max camber 1.1% at 50% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca634221-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca634221-il | 50,000 | 9 | 4.8 at α=-3° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca634221-il | 50,000 | 5 | 10.8 at α=14.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca634221-il | 100,000 | 9 | 28.9 at α=12° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca634221-il | 100,000 | 5 | 34 at α=8° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca634221-il | 200,000 | 9 | 58.1 at α=9.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca634221-il | 200,000 | 5 | 58.4 at α=8.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca634221-il | 500,000 | 9 | 88 at α=8.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca634221-il | 500,000 | 5 | 82 at α=7° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca634221-il | 1,000,000 | 9 | 108 at α=7.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca634221-il | 1,000,000 | 5 | 95.8 at α=6.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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