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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(n6409-il) NACA6409 9% | NACA 6409 Max thickness 9% at 29.3% chord Max camber 6% at 39.6% chord | Remove Airfoil details Airfoil plotter |
(n64012-il) NASA/LANGLEY 64-012 AIRFOIL | NACA 64(1)-012 airfoil Max thickness 12% at 40% chord Max camber 0% at 0% chord | Remove Airfoil details Airfoil plotter |
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Polars for (n6409-il,n64012-il)
Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
---|---|---|---|---|---|---|---|
n6409-il | 50,000 | 9 | 27.1 at α=10.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
n6409-il | 50,000 | 5 | 36 at α=9° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
n6409-il | 100,000 | 9 | 61.6 at α=8.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
n6409-il | 100,000 | 5 | 63.7 at α=6° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
n6409-il | 200,000 | 9 | 87.1 at α=7° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
n6409-il | 200,000 | 5 | 87.4 at α=6° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
n6409-il | 500,000 | 9 | 122.4 at α=5.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
n6409-il | 500,000 | 5 | 118.6 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
n6409-il | 1,000,000 | 9 | 151 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
n6409-il | 1,000,000 | 5 | 144.3 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
n64012-il | 50,000 | 9 | 26.7 at α=5.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
n64012-il | 50,000 | 5 | 27.1 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
n64012-il | 100,000 | 9 | 39.9 at α=4.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
n64012-il | 100,000 | 5 | 32.8 at α=4.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
n64012-il | 200,000 | 9 | 43.5 at α=4.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
n64012-il | 200,000 | 5 | 39.4 at α=4° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
n64012-il | 500,000 | 9 | 56.5 at α=4° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
n64012-il | 500,000 | 5 | 51.5 at α=7.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
n64012-il | 1,000,000 | 9 | 60.9 at α=7.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
n64012-il | 1,000,000 | 5 | 65.4 at α=8° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
Reynolds number calculator |