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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(naca001064-il) NACA 0010-64 | NACA 0010-64 airfoil Max thickness 10% at 40% chord Max camber 0% at 0% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca001064-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca001064-il | 50,000 | 9 | 25.8 at α=4° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca001064-il | 50,000 | 5 | 25.2 at α=3.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca001064-il | 100,000 | 9 | 36.2 at α=3.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca001064-il | 100,000 | 5 | 31.7 at α=3° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca001064-il | 200,000 | 9 | 43.3 at α=3° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca001064-il | 200,000 | 5 | 38.7 at α=5.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca001064-il | 500,000 | 9 | 49.2 at α=5.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca001064-il | 500,000 | 5 | 53.7 at α=9° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca001064-il | 1,000,000 | 9 | 62.5 at α=8.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca001064-il | 1,000,000 | 5 | 71.6 at α=9.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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