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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(naca652215-il) NACA 65(2)-215 | NACA 65(2)-215 airfoil Max thickness 15% at 40% chord Max camber 1.1% at 50% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca652215-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca652215-il | 50,000 | 9 | 22.7 at α=8° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca652215-il | 50,000 | 5 | 23.5 at α=7.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca652215-il | 100,000 | 9 | 46.1 at α=6° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca652215-il | 100,000 | 5 | 39.8 at α=6.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca652215-il | 200,000 | 9 | 57.5 at α=6° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca652215-il | 200,000 | 5 | 55.9 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca652215-il | 500,000 | 9 | 82.5 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca652215-il | 500,000 | 5 | 71.4 at α=4° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca652215-il | 1,000,000 | 9 | 96.4 at α=4.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca652215-il | 1,000,000 | 5 | 79 at α=3.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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