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 | ||
(ultimate-il) ULTIMATE | Carl Goldberg R/C Ultimate Bipe airfoil Max thickness 12.8% at 34.2% chord Max camber 0% at 0% chord | Remove Airfoil details Airfoil plotter |
Drawing Options
Polars for (ultimate-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| ultimate-il | 50,000 | 9 | 28.4 at α=6° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ultimate-il | 50,000 | 5 | 27.5 at α=5.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ultimate-il | 100,000 | 9 | 39 at α=5.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ultimate-il | 100,000 | 5 | 36.1 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ultimate-il | 200,000 | 9 | 47.8 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ultimate-il | 200,000 | 5 | 42 at α=4.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ultimate-il | 500,000 | 9 | 55.4 at α=4.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ultimate-il | 500,000 | 5 | 53.5 at α=7.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ultimate-il | 1,000,000 | 9 | 63.4 at α=7.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ultimate-il | 1,000,000 | 5 | 67 at α=8° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| Reynolds number calculator | |||||||