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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(m21-il) NACA M21 AIRFOIL | NACA/Munk M-21 airfoil Max thickness 12% at 30% chord Max camber 6% at 30% chord | Remove Airfoil details Airfoil plotter |
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Polars for (m21-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| m21-il | 50,000 | 9 | 5.4 at α=8.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m21-il | 50,000 | 5 | 22.2 at α=3.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| m21-il | 100,000 | 9 | 29.8 at α=4° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m21-il | 100,000 | 5 | 48.5 at α=7.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| m21-il | 200,000 | 9 | 70.1 at α=10.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m21-il | 200,000 | 5 | 73.7 at α=7.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| m21-il | 500,000 | 9 | 107.3 at α=8.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m21-il | 500,000 | 5 | 105.8 at α=7.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| m21-il | 1,000,000 | 9 | 135.9 at α=7.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m21-il | 1,000,000 | 5 | 128.4 at α=7.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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