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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(naca001034a08cli02-il) NACA 0010-34 a=0.8 c(li)=0.2 | NACA 0010-34 a=0.8 c(li)=0.2 airfoil Max thickness 10% at 40% chord Max camber 1.3% at 50% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca001034a08cli02-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca001034a08cli02-il | 50,000 | 9 | 26.5 at α=5.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca001034a08cli02-il | 50,000 | 5 | 30.8 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca001034a08cli02-il | 100,000 | 9 | 47.3 at α=4.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca001034a08cli02-il | 100,000 | 5 | 46 at α=4.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca001034a08cli02-il | 200,000 | 9 | 67.1 at α=4° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca001034a08cli02-il | 200,000 | 5 | 62.8 at α=3.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca001034a08cli02-il | 500,000 | 9 | 89.6 at α=3° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca001034a08cli02-il | 500,000 | 5 | 73 at α=2.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca001034a08cli02-il | 1,000,000 | 9 | 93.6 at α=2.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca001034a08cli02-il | 1,000,000 | 5 | 78.7 at α=2° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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