Airfoil Comparison
Select airfoils from the airfoil database or add your own airfoils and compare the airfoil shape and lift/drag polars.
Open full size plan in new window | Open paginated plan in new window | |
Download PDF file | SVG image as text file | |
Clear all | ||
(naca632615-il) NACA 63(2)-615 | NACA 63(2)-615 airfoil Max thickness 15% at 34.8% chord Max camber 3.3% at 50% chord | Remove Airfoil details Airfoil plotter |
Drawing Options
Polars for (naca632615-il)
Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
---|---|---|---|---|---|---|---|
naca632615-il | 50,000 | 9 | 27.8 at α=10.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
naca632615-il | 50,000 | 5 | 24.9 at α=9.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
naca632615-il | 100,000 | 9 | 55 at α=8.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
naca632615-il | 100,000 | 5 | 55.6 at α=7.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
naca632615-il | 200,000 | 9 | 81.7 at α=7.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
naca632615-il | 200,000 | 5 | 78.1 at α=6.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
naca632615-il | 500,000 | 9 | 112.2 at α=6.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
naca632615-il | 500,000 | 5 | 104.5 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
naca632615-il | 1,000,000 | 9 | 134.4 at α=5.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
naca632615-il | 1,000,000 | 5 | 121.6 at α=4° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
Reynolds number calculator |