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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(nlf0115-il) NLF(1)-0115 | NASA/Langley/Somers-Selig-Maughmer NLF(1)-0115 natural laminar flow airfoil Max thickness 15% at 44.1% chord Max camber 1.8% at 30% chord | Remove Airfoil details Airfoil plotter |
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Polars for (nlf0115-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| nlf0115-il | 50,000 | 9 | 30.7 at α=8° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| nlf0115-il | 50,000 | 5 | 26.8 at α=7.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| nlf0115-il | 100,000 | 9 | 45.5 at α=6.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| nlf0115-il | 100,000 | 5 | 41.4 at α=6.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| nlf0115-il | 200,000 | 9 | 60.8 at α=6° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| nlf0115-il | 200,000 | 5 | 56 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| nlf0115-il | 500,000 | 9 | 81.9 at α=4.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| nlf0115-il | 500,000 | 5 | 69.5 at α=3.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| nlf0115-il | 1,000,000 | 9 | 94.4 at α=4° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| nlf0115-il | 1,000,000 | 5 | 76.3 at α=7.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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