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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(naca65206-il) NACA 65-206 | NACA 65-206 airfoil Max thickness 6% at 40% chord Max camber 1.1% at 50% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca65206-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca65206-il | 50,000 | 9 | 22.2 at α=4° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca65206-il | 50,000 | 5 | 29.4 at α=3.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca65206-il | 100,000 | 9 | 40.2 at α=2.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca65206-il | 100,000 | 5 | 39.4 at α=2.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca65206-il | 200,000 | 9 | 54.2 at α=2.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca65206-il | 200,000 | 5 | 50.3 at α=2° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca65206-il | 500,000 | 9 | 71.2 at α=2° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca65206-il | 500,000 | 5 | 56.7 at α=1.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca65206-il | 1,000,000 | 9 | 77.8 at α=1.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca65206-il | 1,000,000 | 5 | 67.7 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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