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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(ag16-il) AG16 | Drela AG16 airfoil Max thickness 7.1% at 23.4% chord Max camber 1.9% at 46.9% chord | Remove Airfoil details Airfoil plotter |
(ls421-il) NASA/LANGLEY LS(1)-0421 AIRFOIL | NASA/Langley LS(1)-0421 general aviation airfoil Max thickness 20.9% at 40% chord Max camber 2.4% at 65% chord | Remove Airfoil details Airfoil plotter |
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Polars for (ag16-il,ls421-il)
Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
---|---|---|---|---|---|---|---|
ag16-il | 50,000 | 9 | 33.5 at α=4.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
ag16-il | 50,000 | 5 | 36 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
ag16-il | 100,000 | 9 | 48.6 at α=4.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
ag16-il | 100,000 | 5 | 48.4 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
ag16-il | 200,000 | 9 | 63.6 at α=4.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
ag16-il | 200,000 | 5 | 60.3 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
ag16-il | 500,000 | 9 | 82.4 at α=4° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
ag16-il | 500,000 | 5 | 76.3 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
ag16-il | 1,000,000 | 9 | 95.3 at α=4.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
ag16-il | 1,000,000 | 5 | 88.4 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
ls421-il | 50,000 | 9 | 23.8 at α=10.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
ls421-il | 50,000 | 5 | 18.4 at α=10° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
ls421-il | 100,000 | 9 | 33.1 at α=8.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
ls421-il | 100,000 | 5 | 30.7 at α=7° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
ls421-il | 200,000 | 9 | 52.4 at α=6.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
ls421-il | 200,000 | 5 | 56.4 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
ls421-il | 500,000 | 9 | 89.5 at α=4.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
ls421-il | 500,000 | 5 | 81.5 at α=3° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
ls421-il | 1,000,000 | 9 | 111.9 at α=3.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
ls421-il | 1,000,000 | 5 | 91.2 at α=1.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
Reynolds number calculator |