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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(n6h10-il) NACA 6-H-10 AIRFOIL | NACA 6-H-10 rotorcraft airfoil Max thickness 10% at 40% chord Max camber 4.6% at 40% chord | Remove Airfoil details Airfoil plotter |
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Polars for (n6h10-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| n6h10-il | 50,000 | 9 | 23.5 at α=10.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n6h10-il | 50,000 | 5 | 17.7 at α=9.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n6h10-il | 100,000 | 9 | 48.3 at α=7.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n6h10-il | 100,000 | 5 | 44.1 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n6h10-il | 200,000 | 9 | 71.5 at α=6.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n6h10-il | 200,000 | 5 | 68.2 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n6h10-il | 500,000 | 9 | 87 at α=6.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n6h10-il | 500,000 | 5 | 87.9 at α=5.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n6h10-il | 1,000,000 | 9 | 107.3 at α=6° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n6h10-il | 1,000,000 | 5 | 99.7 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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