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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(p51htip-il) NACA 66 | NACA 66 Max thickness 12% at 45% chord Max camber 1.3% at 47.5% chord | Remove Airfoil details Airfoil plotter |
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Polars for (p51htip-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| p51htip-il | 50,000 | 9 | 22.2 at α=7° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| p51htip-il | 50,000 | 5 | 22.7 at α=6° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| p51htip-il | 100,000 | 9 | 35.4 at α=4.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| p51htip-il | 100,000 | 5 | 38.7 at α=4.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| p51htip-il | 200,000 | 9 | 52.5 at α=4.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| p51htip-il | 200,000 | 5 | 48 at α=4° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| p51htip-il | 500,000 | 9 | 72.9 at α=3.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| p51htip-il | 500,000 | 5 | 67 at α=3° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| p51htip-il | 1,000,000 | 9 | 90.8 at α=3.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| p51htip-il | 1,000,000 | 5 | 75.1 at α=2.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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