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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(n9-il) N-9 | N-9 airfoil Max thickness 8.5% at 30% chord Max camber 2.7% at 40% chord | Remove Airfoil details Airfoil plotter |
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Polars for (n9-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| n9-il | 50,000 | 9 | 36.8 at α=4.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n9-il | 50,000 | 5 | 38.3 at α=4.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n9-il | 100,000 | 9 | 53.3 at α=4° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n9-il | 100,000 | 5 | 51.6 at α=3.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n9-il | 200,000 | 9 | 68.1 at α=2.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n9-il | 200,000 | 5 | 63.4 at α=3.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n9-il | 500,000 | 9 | 84 at α=1.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n9-il | 500,000 | 5 | 78.3 at α=2.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n9-il | 1,000,000 | 9 | 96.6 at α=2.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n9-il | 1,000,000 | 5 | 92.4 at α=5.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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