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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(naca001065-il) NACA 0010-65 | NACA 0010-65 airfoil Max thickness 10% at 50% chord Max camber 0% at 0% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca001065-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca001065-il | 50,000 | 9 | 20.4 at α=4.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca001065-il | 50,000 | 5 | 21.8 at α=3.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca001065-il | 100,000 | 9 | 27.1 at α=3° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca001065-il | 100,000 | 5 | 28.3 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca001065-il | 200,000 | 9 | 40.6 at α=2.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca001065-il | 200,000 | 5 | 35.7 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca001065-il | 500,000 | 9 | 43.1 at α=6° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca001065-il | 500,000 | 5 | 50.7 at α=10° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca001065-il | 1,000,000 | 9 | 57.4 at α=11° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca001065-il | 1,000,000 | 5 | 65.3 at α=10° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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