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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(m20-il) NACA M20 AIRFOIL | NACA/Munk M-20 airfoil Max thickness 8.2% at 30% chord Max camber 6.1% at 30% chord | Remove Airfoil details Airfoil plotter |
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Polars for (m20-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| m20-il | 50,000 | 9 | 10.4 at α=2.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m20-il | 50,000 | 5 | 26.7 at α=4.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| m20-il | 100,000 | 9 | 55.9 at α=8.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m20-il | 100,000 | 5 | 56.3 at α=8° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| m20-il | 200,000 | 9 | 78.6 at α=7.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m20-il | 200,000 | 5 | 78 at α=7° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| m20-il | 500,000 | 9 | 108.7 at α=6.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m20-il | 500,000 | 5 | 105.2 at α=6.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| m20-il | 1,000,000 | 9 | 131.9 at α=6° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m20-il | 1,000,000 | 5 | 125.3 at α=5.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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