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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(davis-il) Davis AIRFOIL | Consolidated/Davis airfoil (B-24 wing?) Max thickness 11.3% at 30% chord Max camber 3.7% at 40% chord | Remove Airfoil details Airfoil plotter |
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Polars for (davis-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| davis-il | 50,000 | 9 | 22.2 at α=11° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| davis-il | 50,000 | 5 | 36.4 at α=7.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| davis-il | 100,000 | 9 | 54 at α=5.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| davis-il | 100,000 | 5 | 55 at α=5.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| davis-il | 200,000 | 9 | 76.1 at α=4.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| davis-il | 200,000 | 5 | 72.1 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| davis-il | 500,000 | 9 | 101.4 at α=4.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| davis-il | 500,000 | 5 | 91.3 at α=3.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| davis-il | 1,000,000 | 9 | 120.8 at α=3.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| davis-il | 1,000,000 | 5 | 107 at α=5.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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