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 | ||
(s1221-il) S1221 w/o flap | Selig S1221 airfoil Max thickness 12.1% at 21.8% chord Max camber 4.9% at 49.1% chord | Remove Airfoil details Airfoil plotter |
Drawing Options
Polars for (s1221-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| s1221-il | 50,000 | 9 | 27.1 at α=3.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s1221-il | 50,000 | 5 | 36.6 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s1221-il | 100,000 | 9 | 45.2 at α=3.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s1221-il | 100,000 | 5 | 54.5 at α=6° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s1221-il | 200,000 | 9 | 70 at α=11.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s1221-il | 200,000 | 5 | 76.5 at α=7.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s1221-il | 500,000 | 9 | 107.5 at α=9.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s1221-il | 500,000 | 5 | 110 at α=8.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s1221-il | 1,000,000 | 9 | 137.9 at α=8.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s1221-il | 1,000,000 | 5 | 135.1 at α=8° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| Reynolds number calculator | |||||||