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 | ||
(ls013-il) NASA/LANGLEY LS(1)-0013 AIRFOIL | NASA/Langley LS(1)-0013 general aviation airfoil Max thickness 12.9% at 40% chord Max camber 0% at 0% chord | Remove Airfoil details Airfoil plotter |
Drawing Options
Polars for (ls013-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| ls013-il | 50,000 | 9 | 26.4 at α=5.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ls013-il | 50,000 | 5 | 25.9 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ls013-il | 100,000 | 9 | 36.6 at α=4.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ls013-il | 100,000 | 5 | 30.9 at α=6.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ls013-il | 200,000 | 9 | 39.8 at α=4.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ls013-il | 200,000 | 5 | 39.9 at α=7.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ls013-il | 500,000 | 9 | 52.7 at α=8° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ls013-il | 500,000 | 5 | 57.2 at α=8.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ls013-il | 1,000,000 | 9 | 69.2 at α=8.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ls013-il | 1,000,000 | 5 | 71.7 at α=9.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| Reynolds number calculator | |||||||