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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(naca4421-il) NACA 4421 | NACA 4421 airfoil Max thickness 21% at 30% chord Max camber 4% at 40% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca4421-il)
Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
---|---|---|---|---|---|---|---|
naca4421-il | 50,000 | 9 | 3.6 at α=10.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
naca4421-il | 50,000 | 5 | 15.3 at α=2.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
naca4421-il | 100,000 | 9 | 36.1 at α=9.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
naca4421-il | 100,000 | 5 | 41.9 at α=5.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
naca4421-il | 200,000 | 9 | 61.2 at α=7.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
naca4421-il | 200,000 | 5 | 58.4 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
naca4421-il | 500,000 | 9 | 83.3 at α=5.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
naca4421-il | 500,000 | 5 | 78.1 at α=6° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
naca4421-il | 1,000,000 | 9 | 103.1 at α=5.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
naca4421-il | 1,000,000 | 5 | 93.2 at α=5.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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