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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(naca16006-il) NACA 16-006 | NACA 16-006 airfoil Max thickness 6% at 50% chord Max camber 0% at 0% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca16006-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca16006-il | 50,000 | 9 | 16.4 at α=3.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca16006-il | 50,000 | 5 | 16.4 at α=3.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca16006-il | 100,000 | 9 | 20.6 at α=2.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca16006-il | 100,000 | 5 | 21.7 at α=2.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca16006-il | 200,000 | 9 | 25.6 at α=2° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca16006-il | 200,000 | 5 | 33.4 at α=1.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca16006-il | 500,000 | 9 | 41.2 at α=1.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca16006-il | 500,000 | 5 | 34 at α=3.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca16006-il | 1,000,000 | 9 | 40.8 at α=3.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca16006-il | 1,000,000 | 5 | 36.2 at α=3.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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