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 | ||
(usa32-il) USA 32 AIRFOIL | USA-32 airfoil Max thickness 14.7% at 19.8% chord Max camber 9.3% at 39.8% chord | Remove Airfoil details Airfoil plotter |
(usa26-il) USA 26 AIRFOIL | USA-26 airfoil Max thickness 9.8% at 19.9% chord Max camber 4.3% at 39.9% chord | Remove Airfoil details Airfoil plotter |
(us1000root-il) US1000ROOT | Great Planes R/C Ultra-Sport 1000 airfoil Max thickness 18.6% at 27.1% chord Max camber 0% at 0% chord | Remove Airfoil details Airfoil plotter |
Drawing Options
Polars for (usa32-il,usa26-il,us1000root-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| usa32-il | 50,000 | 9 | 4.5 at α=9.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| usa32-il | 50,000 | 5 | 13.5 at α=4° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| usa32-il | 100,000 | 9 | 39.8 at α=5.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| usa32-il | 100,000 | 5 | 41 at α=7.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| usa32-il | 200,000 | 9 | 64.3 at α=7.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| usa32-il | 200,000 | 5 | 50.5 at α=5.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| usa32-il | 500,000 | 9 | 68.9 at α=3.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| usa32-il | 500,000 | 5 | 59.6 at α=3.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| usa32-il | 1,000,000 | 9 | 76.9 at α=3.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| usa32-il | 1,000,000 | 5 | 67.9 at α=5.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| usa26-il | 50,000 | 9 | 27.6 at α=3.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| usa26-il | 50,000 | 5 | 35.3 at α=5.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| usa26-il | 100,000 | 9 | 49.8 at α=9.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| usa26-il | 100,000 | 5 | 52.3 at α=8° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| usa26-il | 200,000 | 9 | 69.5 at α=8.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| usa26-il | 200,000 | 5 | 69.1 at α=7.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| usa26-il | 500,000 | 9 | 95.9 at α=7° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| usa26-il | 500,000 | 5 | 92.9 at α=6° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| usa26-il | 1,000,000 | 9 | 117.6 at α=6.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| usa26-il | 1,000,000 | 5 | 108.7 at α=5.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| us1000root-il | 50,000 | 9 | 19 at α=3.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| us1000root-il | 50,000 | 5 | 24.7 at α=6.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| us1000root-il | 100,000 | 9 | 38.3 at α=8.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| us1000root-il | 100,000 | 5 | 38.6 at α=8° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| us1000root-il | 200,000 | 9 | 49.1 at α=7.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| us1000root-il | 200,000 | 5 | 44.3 at α=7.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| us1000root-il | 500,000 | 9 | 49.8 at α=6.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| us1000root-il | 500,000 | 5 | 43.1 at α=9.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| us1000root-il | 1,000,000 | 9 | 52.3 at α=9° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| us1000root-il | 1,000,000 | 5 | 45.9 at α=8.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| Reynolds number calculator | |||||||