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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(n22-il) N-22 | N-22 airfoil Max thickness 12.4% at 30% chord Max camber 4% at 40% chord | Remove Airfoil details Airfoil plotter |
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Polars for (n22-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| n22-il | 50,000 | 9 | 24.3 at α=7.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n22-il | 50,000 | 5 | 36 at α=5.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n22-il | 100,000 | 9 | 51.8 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n22-il | 100,000 | 5 | 54.1 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n22-il | 200,000 | 9 | 71.7 at α=4.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n22-il | 200,000 | 5 | 70.7 at α=4.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n22-il | 500,000 | 9 | 98.4 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n22-il | 500,000 | 5 | 94.6 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n22-il | 1,000,000 | 9 | 120.5 at α=4.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n22-il | 1,000,000 | 5 | 112.8 at α=4.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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