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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(ls417mod-il) NASA/LANGLEY LS(1)-0417MOD AIRFOIL | NASA/Langley LS(1)-0417MOD general aviation airfoil Max thickness 17% at 30.2% chord Max camber 2.3% at 20.2% chord | Remove Airfoil details Airfoil plotter |
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Polars for (ls417mod-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| ls417mod-il | 50,000 | 9 | 24.3 at α=3.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ls417mod-il | 50,000 | 5 | 26.3 at α=7.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ls417mod-il | 100,000 | 9 | 33.1 at α=8° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ls417mod-il | 100,000 | 5 | 41.1 at α=9° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ls417mod-il | 200,000 | 9 | 52.5 at α=10.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ls417mod-il | 200,000 | 5 | 57.3 at α=8.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ls417mod-il | 500,000 | 9 | 79.9 at α=9° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ls417mod-il | 500,000 | 5 | 81.6 at α=8° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ls417mod-il | 1,000,000 | 9 | 103 at α=8.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ls417mod-il | 1,000,000 | 5 | 100.4 at α=8° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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