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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(n63212-il) NACA 63-212 AIRFOIL | NACA 63(1)-212 airfoil Max thickness 12% at 34.9% chord Max camber 1.1% at 55% chord | Remove Airfoil details Airfoil plotter |
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Polars for (n63212-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| n63212-il | 50,000 | 9 | 33.5 at α=6.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n63212-il | 50,000 | 5 | 32.8 at α=6.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n63212-il | 100,000 | 9 | 49.6 at α=6° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n63212-il | 100,000 | 5 | 46.9 at α=5.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n63212-il | 200,000 | 9 | 65.6 at α=5.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n63212-il | 200,000 | 5 | 57.4 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n63212-il | 500,000 | 9 | 81.7 at α=4.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n63212-il | 500,000 | 5 | 69.1 at α=3.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n63212-il | 1,000,000 | 9 | 91.9 at α=3.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n63212-il | 1,000,000 | 5 | 74.9 at α=7.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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