NACA 65-210 (naca65210-il)
NACA 65-210 - NACA 65-210 airfoil
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(naca65210-il) NACA 65-210 NACA 65-210 airfoil Max thickness 10% 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 65-210 1.00000 0.00000 0.95014 0.00622 0.90028 0.01327 0.85038 0.02057 0.80044 0.02783 0.75045 0.03479 0.70043 0.04128 0.65036 0.04712 0.60027 0.05217 0.55014 0.05625 0.50000 0.05915 0.44984 0.06058 0.39968 0.06067 0.34951 0.05954 0.29936 0.05732 0.24921 0.05397 0.19909 0.04938 0.14899 0.04338 0.09894 0.03555 0.07394 0.03069 0.04898 0.02491 0.02408 0.01757 0.01169 0.01273 0.00678 0.00999 0.00435 0.00819 0.00000 0.00000 0.00565 -0.00719 0.00822 -0.00859 0.01331 -0.01059 0.02592 -0.01385 0.05102 -0.01859 0.07606 -0.02221 0.10106 -0.02521 0.15101 -0.02992 0.20091 -0.03346 0.25079 -0.03607 0.30064 -0.03788 0.35049 -0.03894 0.40032 -0.03925 0.45016 -0.03868 0.50000 -0.03709 0.54986 -0.03435 0.59973 -0.03075 0.64964 -0.02652 0.69957 -0.02184 0.74955 -0.01689 0.79956 -0.01191 0.84962 -0.00711 0.89972 -0.00293 0.94986 0.00010 1.00000 0.00000 |
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Polars for NACA 65-210 (naca65210-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca65210-il | 50,000 | 9 | 24.6 at α=5.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca65210-il | 50,000 | 5 | 29.7 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca65210-il | 100,000 | 9 | 43.3 at α=4.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca65210-il | 100,000 | 5 | 43 at α=3.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca65210-il | 200,000 | 9 | 59 at α=3.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca65210-il | 200,000 | 5 | 50.8 at α=3.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca65210-il | 500,000 | 9 | 72.9 at α=3.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca65210-il | 500,000 | 5 | 62.7 at α=2.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca65210-il | 1,000,000 | 9 | 82.6 at α=2.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca65210-il | 1,000,000 | 5 | 69.2 at α=6° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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