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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(isa962-il) I.S.A. 962 | I.S.A. 962 airfoil Max thickness 9.6% at 20% chord Max camber 4.6% at 40% chord | Remove Airfoil details Airfoil plotter |
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Polars for (isa962-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| isa962-il | 50,000 | 9 | 36.2 at α=5.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| isa962-il | 50,000 | 5 | 39.4 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| isa962-il | 100,000 | 9 | 57.5 at α=4.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| isa962-il | 100,000 | 5 | 56.1 at α=3.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| isa962-il | 200,000 | 9 | 77.1 at α=3° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| isa962-il | 200,000 | 5 | 69.8 at α=2° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| isa962-il | 500,000 | 9 | 98.7 at α=1.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| isa962-il | 500,000 | 5 | 88.3 at α=3.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| isa962-il | 1,000,000 | 9 | 111.8 at α=3.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| isa962-il | 1,000,000 | 5 | 102.5 at α=6.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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