NACA 65-206 (naca65206-il)
NACA 65-206 - NACA 65-206 airfoil
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(naca65206-il) NACA 65-206 NACA 65-206 airfoil Max thickness 6% at 40% chord. Max camber 1.1% at 50% chord Source UIUC Airfoil Coordinates Database Source dat file The dat file is in Selig format |
NACA 65-206 1.00000 0.00000 0.95009 0.00511 0.90018 0.01027 0.85024 0.01538 0.80027 0.02029 0.75028 0.02489 0.70026 0.02907 0.65022 0.03276 0.60016 0.03589 0.55009 0.03836 0.50000 0.04003 0.44990 0.04078 0.39981 0.04069 0.34971 0.03982 0.29962 0.03824 0.24953 0.03592 0.19945 0.03277 0.14939 0.02869 0.09936 0.02340 0.07437 0.02012 0.04939 0.01625 0.02444 0.01140 0.01200 0.00822 0.00706 0.00642 0.00460 0.00524 0.00000 0.00000 0.00540 -0.00424 0.00794 -0.00502 0.01300 -0.00608 0.02556 -0.00768 0.05061 -0.00993 0.07563 -0.01164 0.10064 -0.01306 0.15061 -0.01523 0.20055 -0.01685 0.25047 -0.01802 0.30038 -0.01880 0.35029 -0.01922 0.40019 -0.01927 0.45010 -0.01888 0.50000 -0.01797 0.54991 -0.01646 0.59984 -0.01447 0.64978 -0.01216 0.69974 -0.00963 0.74972 -0.00699 0.79973 -0.00437 0.84976 -0.00192 0.89982 0.00007 0.94991 0.00121 1.00000 0.00000 |
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Polars for NACA 65-206 (naca65206-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca65206-il | 50,000 | 9 | 22.2 at α=4° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca65206-il | 50,000 | 5 | 29.4 at α=3.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca65206-il | 100,000 | 9 | 40.2 at α=2.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca65206-il | 100,000 | 5 | 39.4 at α=2.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca65206-il | 200,000 | 9 | 54.2 at α=2.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca65206-il | 200,000 | 5 | 50.3 at α=2° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca65206-il | 500,000 | 9 | 71.2 at α=2° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca65206-il | 500,000 | 5 | 56.7 at α=1.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca65206-il | 1,000,000 | 9 | 77.8 at α=1.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca65206-il | 1,000,000 | 5 | 67.7 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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