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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(naca671215-il) NACA 67 | NACA 67 Max thickness 15% at 50% chord Max camber 1.1% at 50% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca671215-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca671215-il | 50,000 | 9 | 23.2 at α=7.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca671215-il | 50,000 | 5 | 21.4 at α=6.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca671215-il | 100,000 | 9 | 34.9 at α=5.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca671215-il | 100,000 | 5 | 25.8 at α=6.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca671215-il | 200,000 | 9 | 42.4 at α=5.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca671215-il | 200,000 | 5 | 33.4 at α=5.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca671215-il | 500,000 | 9 | 56.2 at α=4.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca671215-il | 500,000 | 5 | 55.5 at α=3.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca671215-il | 1,000,000 | 9 | 84.3 at α=3.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca671215-il | 1,000,000 | 5 | 71.3 at α=2.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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