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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(naca63a210-il) NACA 63A210 | NACA 63A210 airfoil Max thickness 10% at 34.9% chord Max camber 1.3% at 50% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca63a210-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca63a210-il | 50,000 | 9 | 28.3 at α=6.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca63a210-il | 50,000 | 5 | 31.3 at α=4.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca63a210-il | 100,000 | 9 | 37.6 at α=6.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca63a210-il | 100,000 | 5 | 40.2 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca63a210-il | 200,000 | 9 | 44.9 at α=6° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca63a210-il | 200,000 | 5 | 48.3 at α=4.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca63a210-il | 500,000 | 9 | 52 at α=6.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca63a210-il | 500,000 | 5 | 58.3 at α=6° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca63a210-il | 1,000,000 | 9 | 58.5 at α=7.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca63a210-il | 1,000,000 | 5 | 70.4 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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