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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(raf69-il) RAF 69 AIRFOIL | RAF-69 airfoil Max thickness 20.6% at 30% chord Max camber 1.7% at 50% chord | Remove Airfoil details Airfoil plotter |
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Polars for (raf69-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| raf69-il | 50,000 | 9 | 4.7 at α=2.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| raf69-il | 50,000 | 5 | 18.7 at α=5.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| raf69-il | 100,000 | 9 | 40.3 at α=8° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| raf69-il | 100,000 | 5 | 40.1 at α=7.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| raf69-il | 200,000 | 9 | 58.2 at α=8° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| raf69-il | 200,000 | 5 | 51.7 at α=8° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| raf69-il | 500,000 | 9 | 74 at α=8.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| raf69-il | 500,000 | 5 | 63.9 at α=8.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| raf69-il | 1,000,000 | 9 | 88.1 at α=8.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| raf69-il | 1,000,000 | 5 | 72.4 at α=9.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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