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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(naca0024-il) NACA 0024 | NACA 0024 airfoil Max thickness 24% at 30% chord Max camber 0% at 0% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca0024-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca0024-il | 50,000 | 9 | 15.4 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca0024-il | 50,000 | 5 | 21.6 at α=8° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca0024-il | 100,000 | 9 | 34.7 at α=8.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca0024-il | 100,000 | 5 | 30.6 at α=9.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca0024-il | 200,000 | 9 | 43 at α=10° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca0024-il | 200,000 | 5 | 37 at α=9.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca0024-il | 500,000 | 9 | 52.1 at α=10.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca0024-il | 500,000 | 5 | 50.8 at α=8.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca0024-il | 1,000,000 | 9 | 68.8 at α=9° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca0024-il | 1,000,000 | 5 | 64.1 at α=9.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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