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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(p51hroot-il) NACA 66 | NACA 66 Max thickness 15.5% at 45% chord Max camber 1.3% at 47.5% chord | Remove Airfoil details Airfoil plotter |
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Polars for (p51hroot-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| p51hroot-il | 50,000 | 9 | 20.3 at α=9.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| p51hroot-il | 50,000 | 5 | 19.5 at α=7.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| p51hroot-il | 100,000 | 9 | 32.3 at α=5.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| p51hroot-il | 100,000 | 5 | 27.1 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| p51hroot-il | 200,000 | 9 | 42.4 at α=6° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| p51hroot-il | 200,000 | 5 | 43.6 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| p51hroot-il | 500,000 | 9 | 67.2 at α=4° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| p51hroot-il | 500,000 | 5 | 55.6 at α=2.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| p51hroot-il | 1,000,000 | 9 | 70.1 at α=2.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| p51hroot-il | 1,000,000 | 5 | 46.5 at α=7.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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