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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(m16-il) NACA M16 AIRFOIL | NACA/Munk M-16 airfoil Max thickness 6.2% at 30% chord Max camber 4.1% at 30% chord | Remove Airfoil details Airfoil plotter |
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Polars for (m16-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| m16-il | 50,000 | 9 | 39 at α=6.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m16-il | 50,000 | 5 | 38.1 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| m16-il | 100,000 | 9 | 53.6 at α=6° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m16-il | 100,000 | 5 | 53.6 at α=5.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| m16-il | 200,000 | 9 | 70.3 at α=5.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m16-il | 200,000 | 5 | 68.9 at α=5.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| m16-il | 500,000 | 9 | 92.7 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m16-il | 500,000 | 5 | 88.6 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| m16-il | 1,000,000 | 9 | 109.1 at α=4.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m16-il | 1,000,000 | 5 | 102.9 at α=4.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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