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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(naca0008-il) NACA 0008 | NACA 0008 airfoil Max thickness 8% at 30% chord Max camber 0% at 0% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca0008-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca0008-il | 50,000 | 9 | 26.1 at α=4° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca0008-il | 50,000 | 5 | 24.6 at α=3.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca0008-il | 100,000 | 9 | 34.6 at α=3.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca0008-il | 100,000 | 5 | 30.4 at α=3.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca0008-il | 200,000 | 9 | 41.1 at α=3.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca0008-il | 200,000 | 5 | 38.5 at α=5.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca0008-il | 500,000 | 9 | 50.2 at α=5.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca0008-il | 500,000 | 5 | 56.2 at α=7° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca0008-il | 1,000,000 | 9 | 68.3 at α=6.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca0008-il | 1,000,000 | 5 | 69.9 at α=7.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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