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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(naca64208-il) NACA 64-208 | NACA 64-208 airfoil Max thickness 8% at 40% chord Max camber 1.1% at 50% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca64208-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca64208-il | 50,000 | 9 | 26.6 at α=4.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca64208-il | 50,000 | 5 | 31.5 at α=4° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca64208-il | 100,000 | 9 | 45.3 at α=3.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca64208-il | 100,000 | 5 | 42.8 at α=3.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca64208-il | 200,000 | 9 | 58.9 at α=3.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca64208-il | 200,000 | 5 | 53.3 at α=3° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca64208-il | 500,000 | 9 | 75.8 at α=2.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca64208-il | 500,000 | 5 | 60.7 at α=2.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca64208-il | 1,000,000 | 9 | 80.8 at α=2.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca64208-il | 1,000,000 | 5 | 70.1 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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