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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(naca001035-il) NACA 0010-35 | NACA 0010-35 airfoil Max thickness 10% at 50% chord Max camber 0% at 0% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca001035-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca001035-il | 50,000 | 9 | 17.6 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca001035-il | 50,000 | 5 | 18.8 at α=4.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca001035-il | 100,000 | 9 | 24 at α=3.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca001035-il | 100,000 | 5 | 27.1 at α=2.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca001035-il | 200,000 | 9 | 37.4 at α=2.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca001035-il | 200,000 | 5 | 32.2 at α=2.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca001035-il | 500,000 | 9 | 39.6 at α=2.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca001035-il | 500,000 | 5 | 36.9 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca001035-il | 1,000,000 | 9 | 43.2 at α=2.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca001035-il | 1,000,000 | 5 | 45.2 at α=6.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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