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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(naca2411-il) NACA 2.5411 | NACA 2411 airfoil Max thickness 11% at 29.5% chord Max camber 2.5% at 39.6% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca2411-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca2411-il | 50,000 | 9 | 34.3 at α=7.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca2411-il | 50,000 | 5 | 35.9 at α=6.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca2411-il | 100,000 | 9 | 52.9 at α=6.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca2411-il | 100,000 | 5 | 52.2 at α=5.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca2411-il | 200,000 | 9 | 71.2 at α=5.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca2411-il | 200,000 | 5 | 67.4 at α=5.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca2411-il | 500,000 | 9 | 95.6 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca2411-il | 500,000 | 5 | 86.6 at α=4° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca2411-il | 1,000,000 | 9 | 113.7 at α=4° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca2411-il | 1,000,000 | 5 | 97.7 at α=4.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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