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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(rae101-il) RAE 101 AIRFOIL | RAE 101 airfoil Max thickness 10% at 30% chord Max camber 0% at 0% chord | Remove Airfoil details Airfoil plotter |
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Polars for (rae101-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| rae101-il | 50,000 | 9 | 27.9 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| rae101-il | 50,000 | 5 | 27.4 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| rae101-il | 100,000 | 9 | 39 at α=4.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| rae101-il | 100,000 | 5 | 36.4 at α=4.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| rae101-il | 200,000 | 9 | 49.7 at α=4.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| rae101-il | 200,000 | 5 | 44.2 at α=3.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| rae101-il | 500,000 | 9 | 60.3 at α=3.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| rae101-il | 500,000 | 5 | 56.4 at α=6.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| rae101-il | 1,000,000 | 9 | 67.9 at α=5.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| rae101-il | 1,000,000 | 5 | 70.2 at α=7° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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