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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(naca0015-il) NACA 0015 | NACA 0015 airfoil Max thickness 15% at 30% chord Max camber 0% at 0% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca0015-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca0015-il | 50,000 | 9 | 24.7 at α=6.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca0015-il | 50,000 | 5 | 26.9 at α=6.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca0015-il | 100,000 | 9 | 37.5 at α=6° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca0015-il | 100,000 | 5 | 37.8 at α=6.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca0015-il | 200,000 | 9 | 49.6 at α=6.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca0015-il | 200,000 | 5 | 48.6 at α=7° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca0015-il | 500,000 | 9 | 66.4 at α=7.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca0015-il | 500,000 | 5 | 63.2 at α=8.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca0015-il | 1,000,000 | 9 | 77.9 at α=9° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca0015-il | 1,000,000 | 5 | 75.1 at α=10.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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