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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(usa50-il) USA 50 AIRFOIL | USA-50 airfoil Max thickness 7% at 30% chord Max camber 2% at 40% chord | Remove Airfoil details Airfoil plotter |
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Polars for (usa50-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| usa50-il | 50,000 | 9 | 36.4 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| usa50-il | 50,000 | 5 | 36.1 at α=4.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| usa50-il | 100,000 | 9 | 51.1 at α=4° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| usa50-il | 100,000 | 5 | 49.2 at α=3.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| usa50-il | 200,000 | 9 | 65.1 at α=2.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| usa50-il | 200,000 | 5 | 60.5 at α=3° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| usa50-il | 500,000 | 9 | 80.4 at α=2.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| usa50-il | 500,000 | 5 | 73.3 at α=2.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| usa50-il | 1,000,000 | 9 | 92.3 at α=2.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| usa50-il | 1,000,000 | 5 | 83.7 at α=2° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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