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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(naca644421-il) NACA 64(4)-421 | NACA 64(4)-421 airfoil Max thickness 20.9% at 34.8% chord Max camber 2.2% at 50% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca644421-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca644421-il | 50,000 | 9 | 3.9 at α=11.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca644421-il | 50,000 | 5 | 11.3 at α=14.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca644421-il | 100,000 | 9 | 21.9 at α=12.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca644421-il | 100,000 | 5 | 26.1 at α=7.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca644421-il | 200,000 | 9 | 61.5 at α=9.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca644421-il | 200,000 | 5 | 62.6 at α=8° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca644421-il | 500,000 | 9 | 98.5 at α=8.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca644421-il | 500,000 | 5 | 93.2 at α=7.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca644421-il | 1,000,000 | 9 | 124.1 at α=7.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca644421-il | 1,000,000 | 5 | 112.5 at α=6.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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