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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(naca63206-il) NACA 63-206 | NACA 63-206 airfoil Max thickness 6% at 35% chord Max camber 1.1% at 50% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca63206-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca63206-il | 50,000 | 9 | 24.2 at α=4° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca63206-il | 50,000 | 5 | 31.1 at α=3.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca63206-il | 100,000 | 9 | 43.1 at α=3° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca63206-il | 100,000 | 5 | 41.5 at α=2.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca63206-il | 200,000 | 9 | 57 at α=2.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca63206-il | 200,000 | 5 | 52.4 at α=2.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca63206-il | 500,000 | 9 | 72.4 at α=2.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca63206-il | 500,000 | 5 | 58.7 at α=1.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca63206-il | 1,000,000 | 9 | 80.3 at α=2° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca63206-il | 1,000,000 | 5 | 69 at α=5.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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