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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(nlf1015-il) NASA NLF1015 | NASA/Langley/Somers-Maughmer NLF(1)-1015 natural laminar flow airfoil Max thickness 15% at 39.8% chord Max camber 4.3% at 62.8% chord | Remove Airfoil details Airfoil plotter |
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Polars for (nlf1015-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| nlf1015-il | 50,000 | 9 | 25.8 at α=11.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| nlf1015-il | 50,000 | 5 | 21.9 at α=11.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| nlf1015-il | 100,000 | 9 | 40.7 at α=10.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| nlf1015-il | 100,000 | 5 | 38.6 at α=9.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| nlf1015-il | 200,000 | 9 | 75.6 at α=7.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| nlf1015-il | 200,000 | 5 | 77.2 at α=6.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| nlf1015-il | 500,000 | 9 | 130.6 at α=6° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| nlf1015-il | 500,000 | 5 | 125.9 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| nlf1015-il | 1,000,000 | 9 | 167.7 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| nlf1015-il | 1,000,000 | 5 | 154.6 at α=3.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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