Airfoil Comparison
Select airfoils from the airfoil database or add your own airfoils and compare the airfoil shape and lift/drag polars.
| Open full size plan in new window | Open paginated plan in new window | |
| Download PDF file | SVG image as text file | |
| Clear all | ||
(s2050-il) S2050 8.93% | Selig S2050 low Reynolds number airfoil Max thickness 8.9% at 29.8% chord Max camber 1.1% at 69.5% chord | Remove Airfoil details Airfoil plotter |
Drawing Options
Polars for (s2050-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| s2050-il | 50,000 | 9 | 33.2 at α=5.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s2050-il | 50,000 | 5 | 34.6 at α=5.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s2050-il | 100,000 | 9 | 48.8 at α=4.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s2050-il | 100,000 | 5 | 47 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s2050-il | 200,000 | 9 | 63.5 at α=4.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s2050-il | 200,000 | 5 | 58.2 at α=4.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s2050-il | 500,000 | 9 | 80.9 at α=4.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s2050-il | 500,000 | 5 | 70.6 at α=4.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s2050-il | 1,000,000 | 9 | 88.5 at α=3.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s2050-il | 1,000,000 | 5 | 80.8 at α=7° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| Reynolds number calculator | |||||||