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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(s1091-il) S1091 | Selig S1091 airfoil Max thickness 5.1% at 13.4% chord Max camber 5.7% at 45.1% chord | Remove Airfoil details Airfoil plotter |
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Polars for (s1091-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| s1091-il | 50,000 | 9 | 46.1 at α=6.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s1091-il | 50,000 | 5 | 48.3 at α=5.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s1091-il | 100,000 | 9 | 70.6 at α=5.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s1091-il | 100,000 | 5 | 69.9 at α=4.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s1091-il | 200,000 | 9 | 96.2 at α=4° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s1091-il | 200,000 | 5 | 90.7 at α=3.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s1091-il | 500,000 | 9 | 129.8 at α=2.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s1091-il | 500,000 | 5 | 111 at α=2.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s1091-il | 1,000,000 | 9 | 153 at α=2.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s1091-il | 1,000,000 | 5 | 117.7 at α=3.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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