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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(naca654421a05-il) NACA 65(4)-421 a=0.5 | NACA 65(4)-421 a=0.5 airfoil Max thickness 21% at 39.9% chord Max camber 3% at 45% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca654421a05-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca654421a05-il | 50,000 | 9 | 3.7 at α=12.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca654421a05-il | 50,000 | 5 | 10.2 at α=14.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca654421a05-il | 100,000 | 9 | 17.4 at α=13.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca654421a05-il | 100,000 | 5 | 16.5 at α=12.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca654421a05-il | 200,000 | 9 | 31.7 at α=11.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca654421a05-il | 200,000 | 5 | 44.4 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca654421a05-il | 500,000 | 9 | 96.5 at α=7.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca654421a05-il | 500,000 | 5 | 92.5 at α=7.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca654421a05-il | 1,000,000 | 9 | 128.2 at α=7.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca654421a05-il | 1,000,000 | 5 | 117.6 at α=6.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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