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 | ||
(naca0021-il) NACA 0021 | NACA 0021 airfoil Max thickness 21% at 30% chord Max camber 0% at 0% chord | Remove Airfoil details Airfoil plotter |
Drawing Options
Polars for (naca0021-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca0021-il | 50,000 | 9 | 19.1 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca0021-il | 50,000 | 5 | 25.2 at α=7° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca0021-il | 100,000 | 9 | 36.6 at α=7.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca0021-il | 100,000 | 5 | 35.6 at α=8.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca0021-il | 200,000 | 9 | 48 at α=8.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca0021-il | 200,000 | 5 | 43.6 at α=9.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca0021-il | 500,000 | 9 | 60.6 at α=10.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca0021-il | 500,000 | 5 | 56.8 at α=8.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca0021-il | 1,000,000 | 9 | 74.7 at α=8.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca0021-il | 1,000,000 | 5 | 71 at α=9° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| Reynolds number calculator | |||||||