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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(raf19-il) RAF 19 AIRFOIL | RAF-19 airfoil Max thickness 10.5% at 15% chord Max camber 10.1% at 40% chord | Remove Airfoil details Airfoil plotter |
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Polars for (raf19-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| raf19-il | 50,000 | 9 | 13.1 at α=3° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| raf19-il | 50,000 | 5 | 32.2 at α=5.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| raf19-il | 100,000 | 9 | 44.6 at α=6.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| raf19-il | 100,000 | 5 | 48.7 at α=6° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| raf19-il | 200,000 | 9 | 62.4 at α=9.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| raf19-il | 200,000 | 5 | 61.8 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| raf19-il | 500,000 | 9 | 82.2 at α=7° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| raf19-il | 500,000 | 5 | 83.3 at α=6° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| raf19-il | 1,000,000 | 9 | 104.8 at α=6.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| raf19-il | 1,000,000 | 5 | 103 at α=5.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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