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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(oa206-il) ONERA OA206 AIRFOIL | ONERA/Aerospatiale OA206 rotorcraft airfoil (Constructed from patent and smoothed) Max thickness 6% at 31.8% chord Max camber 0.8% at 19.6% chord | Remove Airfoil details Airfoil plotter |
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Polars for (oa206-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| oa206-il | 50,000 | 9 | 27.3 at α=4° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| oa206-il | 50,000 | 5 | 27.1 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| oa206-il | 100,000 | 9 | 36.4 at α=3.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| oa206-il | 100,000 | 5 | 35.9 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| oa206-il | 200,000 | 9 | 46.1 at α=4.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| oa206-il | 200,000 | 5 | 44.1 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| oa206-il | 500,000 | 9 | 58.6 at α=4.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| oa206-il | 500,000 | 5 | 56.3 at α=6° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| oa206-il | 1,000,000 | 9 | 67.6 at α=4.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| oa206-il | 1,000,000 | 5 | 69.7 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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