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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(naca23012-il) NACA 23012 12% | NACA 23012 airfoil Max thickness 12% at 29.8% chord Max camber 1.8% at 12.7% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca23012-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca23012-il | 50,000 | 9 | 24.2 at α=4.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca23012-il | 50,000 | 5 | 28.3 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca23012-il | 100,000 | 9 | 36.4 at α=5.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca23012-il | 100,000 | 5 | 40.5 at α=7.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca23012-il | 200,000 | 9 | 51 at α=9° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca23012-il | 200,000 | 5 | 55.4 at α=8.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca23012-il | 500,000 | 9 | 77.1 at α=9° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca23012-il | 500,000 | 5 | 77.2 at α=8.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca23012-il | 1,000,000 | 9 | 96.8 at α=8.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca23012-il | 1,000,000 | 5 | 93.7 at α=7.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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