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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(ls413mod-il) NASA/LANGLEY LS(1)-0413MOD AIRFOIL | NASA/Langley LS(1)-0413MOD general aviation airfoil Max thickness 13% at 35% chord Max camber 2.2% at 40% chord | Remove Airfoil details Airfoil plotter |
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Polars for (ls413mod-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| ls413mod-il | 50,000 | 9 | 34.9 at α=6° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ls413mod-il | 50,000 | 5 | 35.8 at α=5.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ls413mod-il | 100,000 | 9 | 53.3 at α=5.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ls413mod-il | 100,000 | 5 | 49.2 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ls413mod-il | 200,000 | 9 | 66 at α=4.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ls413mod-il | 200,000 | 5 | 62 at α=3.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ls413mod-il | 500,000 | 9 | 86.7 at α=3° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ls413mod-il | 500,000 | 5 | 77.6 at α=6.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ls413mod-il | 1,000,000 | 9 | 99.8 at α=2.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ls413mod-il | 1,000,000 | 5 | 92.5 at α=6° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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