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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(n0011sc-il) NACA/LANGLEY SYMMETRICAL | NASA/Langley 11% symmetrical supercritical airfoil Max thickness 11% at 40% chord Max camber 0% at 0% chord | Remove Airfoil details Airfoil plotter |
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Polars for (n0011sc-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| n0011sc-il | 50,000 | 9 | 22.3 at α=4° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n0011sc-il | 50,000 | 5 | 22.3 at α=3.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n0011sc-il | 100,000 | 9 | 30.4 at α=3° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n0011sc-il | 100,000 | 5 | 30.5 at α=4.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n0011sc-il | 200,000 | 9 | 41.5 at α=2° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n0011sc-il | 200,000 | 5 | 39.2 at α=7° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n0011sc-il | 500,000 | 9 | 49.6 at α=7° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n0011sc-il | 500,000 | 5 | 50.4 at α=8.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n0011sc-il | 1,000,000 | 9 | 59 at α=8° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n0011sc-il | 1,000,000 | 5 | 62.6 at α=9.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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