NACA 66-206 (naca66206-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 66-206 (naca66206-il) Reynolds number: 100,000 Max Cl/Cd: 28.44 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca66206-il-100000.txt Download as CSV file: xf-naca66206-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 66-206
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.5432 0.09681 0.09191 -0.0216 1.0000 0.0593
-8.250 -0.5508 0.09359 0.08876 -0.0266 1.0000 0.0596
-8.000 -0.5546 0.09050 0.08562 -0.0313 1.0000 0.0599
-7.750 -0.5396 0.08423 0.07949 -0.0240 1.0000 0.0639
-7.500 -0.5369 0.08057 0.07585 -0.0256 1.0000 0.0665
-7.250 -0.5355 0.07674 0.07200 -0.0285 1.0000 0.0696
-7.000 -0.5344 0.07357 0.06868 -0.0336 1.0000 0.0725
-6.500 -0.5231 0.06508 0.06024 -0.0328 1.0000 0.0786
-6.250 -0.5165 0.06334 0.05804 -0.0356 1.0000 0.0863
-6.000 -0.5094 0.05783 0.05276 -0.0340 1.0000 0.0899
-5.750 -0.5002 0.05587 0.05035 -0.0348 1.0000 0.1006
-5.500 -0.4905 0.05147 0.04622 -0.0329 1.0000 0.1068
-5.250 -0.4804 0.04838 0.04300 -0.0322 1.0000 0.1189
-5.000 -0.4694 0.04558 0.04008 -0.0312 1.0000 0.1343
-4.750 -0.4588 0.04282 0.03719 -0.0303 1.0000 0.1568
-4.500 -0.4482 0.04020 0.03449 -0.0287 1.0000 0.1849
-4.250 -0.4388 0.03784 0.03219 -0.0266 1.0000 0.2250
-3.250 -0.3138 0.02514 0.01662 -0.0245 1.0000 0.0826
-3.000 -0.2863 0.02279 0.01376 -0.0228 1.0000 0.0689
-2.750 -0.2599 0.02087 0.01140 -0.0214 1.0000 0.0631
-2.500 -0.2348 0.01942 0.00972 -0.0202 1.0000 0.0618
-2.250 -0.2110 0.01800 0.00831 -0.0192 1.0000 0.0665
-2.000 -0.1875 0.01694 0.00720 -0.0182 1.0000 0.0750
-1.750 -0.1646 0.01585 0.00610 -0.0172 1.0000 0.0814
-1.500 -0.1414 0.01504 0.00523 -0.0163 1.0000 0.0949
-1.250 -0.1098 0.01166 0.00496 -0.0136 1.0000 1.0000
-1.000 -0.0925 0.01168 0.00465 -0.0121 1.0000 1.0000
-0.750 -0.0733 0.01174 0.00446 -0.0111 1.0000 1.0000
-0.500 -0.0529 0.01182 0.00436 -0.0103 1.0000 1.0000
-0.250 -0.0317 0.01193 0.00431 -0.0096 1.0000 1.0000
0.000 -0.0102 0.01206 0.00427 -0.0090 1.0000 1.0000
0.250 0.0117 0.01222 0.00432 -0.0085 1.0000 1.0000
0.500 0.0337 0.01240 0.00442 -0.0080 1.0000 1.0000
0.750 0.0557 0.01259 0.00456 -0.0076 1.0000 1.0000
1.000 0.0778 0.01281 0.00475 -0.0072 1.0000 1.0000
1.250 0.0999 0.01305 0.00499 -0.0068 1.0000 1.0000
1.500 0.1219 0.01332 0.00528 -0.0065 1.0000 1.0000
1.750 0.1438 0.01360 0.00559 -0.0061 1.0000 1.0000
2.000 0.1655 0.01392 0.00596 -0.0058 1.0000 1.0000
2.250 0.1871 0.01426 0.00637 -0.0054 1.0000 1.0000
2.500 0.2084 0.01463 0.00684 -0.0051 1.0000 1.0000
2.750 0.2296 0.01504 0.00737 -0.0048 1.0000 1.0000
3.000 0.2673 0.01570 0.00835 -0.0079 0.9934 1.0000
3.250 0.4692 0.01650 0.00642 -0.0274 0.0806 1.0000
3.500 0.4899 0.01778 0.00766 -0.0260 0.0677 1.0000
3.750 0.5128 0.01982 0.00951 -0.0249 0.0634 1.0000
4.000 0.5410 0.02158 0.01150 -0.0241 0.0622 1.0000
4.250 0.5704 0.02399 0.01408 -0.0235 0.0627 1.0000
4.500 0.5979 0.02584 0.01634 -0.0225 0.0603 1.0000
4.750 0.6246 0.02837 0.01927 -0.0214 0.0602 1.0000
5.000 0.6534 0.03168 0.02321 -0.0196 0.0728 1.0000
6.750 0.7864 0.05953 0.05421 -0.0111 0.1189 1.0000
7.000 0.8005 0.06584 0.06027 -0.0108 0.1031 1.0000
7.250 0.7959 0.06698 0.06208 -0.0102 0.0935 1.0000
7.500 0.8061 0.07309 0.06805 -0.0099 0.0891 1.0000
7.750 0.7954 0.07553 0.07093 -0.0109 0.0829 1.0000
8.000 0.8055 0.07960 0.07497 -0.0101 0.0784 1.0000
8.250 0.8107 0.08607 0.08136 -0.0102 0.0761 1.0000
8.500 0.7937 0.08916 0.08466 -0.0122 0.0753 1.0000
8.750 0.7766 0.09300 0.08855 -0.0146 0.0745 1.0000
9.000 0.7663 0.09790 0.09344 -0.0189 0.0741 1.0000
9.250 0.7576 0.10317 0.09865 -0.0243 0.0725 1.0000
9.500 0.7542 0.10791 0.10334 -0.0276 0.0699 1.0000
9.750 0.7553 0.11225 0.10767 -0.0294 0.0676 1.0000
10.000 0.6609 0.11367 0.10942 -0.0240 0.0755 1.0000
10.250 0.6499 0.11825 0.11396 -0.0280 0.0743 1.0000
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Polar data table (+)
Polar graphs
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