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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(n63010a-il) NACA 63A010 AIRFOIL | NACA 63-010A airfoil Max thickness 10% at 35% chord Max camber 0% at 0% chord | Remove Airfoil details Airfoil plotter |
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Polars for (n63010a-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| n63010a-il | 50,000 | 9 | 27.3 at α=4.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n63010a-il | 50,000 | 5 | 26.4 at α=4.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n63010a-il | 100,000 | 9 | 37.7 at α=4° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n63010a-il | 100,000 | 5 | 34.9 at α=3.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n63010a-il | 200,000 | 9 | 47.7 at α=3.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n63010a-il | 200,000 | 5 | 39.2 at α=3.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n63010a-il | 500,000 | 9 | 52.3 at α=3.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n63010a-il | 500,000 | 5 | 51 at α=6° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n63010a-il | 1,000,000 | 9 | 59.7 at α=5.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n63010a-il | 1,000,000 | 5 | 64.2 at α=7° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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