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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(naca16021-il) NACA 16-021 | NACA 16-021 airfoil Max thickness 21% at 50% chord Max camber 0% at 0% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca16021-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca16021-il | 50,000 | 9 | 21.9 at α=8.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca16021-il | 50,000 | 5 | 17.9 at α=7.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca16021-il | 100,000 | 9 | 28.3 at α=6.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca16021-il | 100,000 | 5 | 20.7 at α=7° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca16021-il | 200,000 | 9 | 33.7 at α=6° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca16021-il | 200,000 | 5 | 26.3 at α=6° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca16021-il | 500,000 | 9 | 39.5 at α=5.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca16021-il | 500,000 | 5 | 28.4 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca16021-il | 1,000,000 | 9 | 42 at α=4.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca16021-il | 1,000,000 | 5 | 31.6 at α=12.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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