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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(ls421mod-il) NASA/LANGLEY LS(1)-0421MOD AIRFOIL | NASA/Langley LS(1)-0421MOD general aviation airfoil Max thickness 21% at 32.5% chord Max camber 2.3% at 30% chord | Remove Airfoil details Airfoil plotter |
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Polars for (ls421mod-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| ls421mod-il | 50,000 | 9 | 7.2 at α=0.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ls421mod-il | 50,000 | 5 | 18.6 at α=7.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ls421mod-il | 100,000 | 9 | 37.2 at α=8.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ls421mod-il | 100,000 | 5 | 38.4 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ls421mod-il | 200,000 | 9 | 51.9 at α=7° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ls421mod-il | 200,000 | 5 | 56 at α=6° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ls421mod-il | 500,000 | 9 | 80.8 at α=6° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ls421mod-il | 500,000 | 5 | 77.1 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ls421mod-il | 1,000,000 | 9 | 100 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ls421mod-il | 1,000,000 | 5 | 90.3 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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