NACA 64-008A AIRFOIL (n64008a-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 64-008A AIRFOIL (n64008a-il) Reynolds number: 100,000 Max Cl/Cd: 32.45 at α=3° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n64008a-il-100000.txt Download as CSV file: xf-n64008a-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 64-008A AIRFOIL
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 -0.6995 0.08631 0.08172 -0.0046 1.0000 0.1120
-8.250 -0.5878 0.07394 0.06960 -0.0071 1.0000 0.1349
-8.000 -0.7019 0.07714 0.07254 -0.0076 1.0000 0.1206
-7.750 -0.7297 0.07381 0.06869 -0.0138 1.0000 0.1276
-7.500 -0.7029 0.06836 0.06362 -0.0108 1.0000 0.1352
-7.250 -0.7020 0.06404 0.05917 -0.0120 1.0000 0.1455
-7.000 -0.6965 0.06023 0.05526 -0.0122 1.0000 0.1587
-6.750 -0.6867 0.05681 0.05179 -0.0117 1.0000 0.1741
-6.500 -0.6806 0.05364 0.04847 -0.0113 1.0000 0.1983
-5.750 -0.6102 0.03502 0.02693 -0.0101 1.0000 0.0725
-5.500 -0.5886 0.03161 0.02314 -0.0090 1.0000 0.0710
-5.250 -0.5648 0.02830 0.01940 -0.0077 1.0000 0.0664
-5.000 -0.5390 0.02571 0.01623 -0.0062 1.0000 0.0633
-4.750 -0.5129 0.02356 0.01384 -0.0052 1.0000 0.0634
-4.500 -0.4873 0.02173 0.01192 -0.0044 1.0000 0.0659
-4.250 -0.4624 0.02054 0.01058 -0.0034 1.0000 0.0728
-4.000 -0.4389 0.01873 0.00888 -0.0023 1.0000 0.0773
-3.750 -0.4172 0.01748 0.00763 -0.0009 1.0000 0.0853
-3.500 -0.3977 0.01620 0.00644 0.0007 1.0000 0.1023
-3.250 -0.3865 0.01327 0.00499 0.0030 1.0000 0.3228
-3.000 -0.3875 0.01194 0.00559 0.0116 1.0000 0.7642
-2.750 -0.3727 0.01215 0.00585 0.0167 1.0000 0.8366
-2.500 -0.3452 0.01294 0.00658 0.0212 1.0000 0.9076
-2.250 -0.2475 0.01397 0.00706 0.0109 1.0000 0.9550
-2.000 -0.1923 0.01375 0.00659 0.0049 1.0000 0.9681
-1.750 -0.1395 0.01341 0.00605 -0.0008 1.0000 0.9794
-1.500 -0.0894 0.01303 0.00554 -0.0063 1.0000 0.9902
-1.250 -0.0418 0.01264 0.00505 -0.0114 1.0000 1.0000
-1.000 -0.0245 0.01233 0.00473 -0.0108 1.0000 1.0000
-0.750 -0.0092 0.01206 0.00447 -0.0097 1.0000 1.0000
-0.500 0.0023 0.01186 0.00428 -0.0080 1.0000 1.0000
-0.250 0.0060 0.01171 0.00418 -0.0048 1.0000 1.0000
0.000 0.0000 0.01166 0.00415 0.0000 1.0000 1.0000
0.250 -0.0059 0.01171 0.00418 0.0048 1.0000 1.0000
0.500 -0.0023 0.01185 0.00428 0.0080 1.0000 1.0000
0.750 0.0092 0.01206 0.00447 0.0097 1.0000 1.0000
1.000 0.0246 0.01232 0.00473 0.0108 1.0000 1.0000
1.250 0.0418 0.01264 0.00505 0.0114 1.0000 1.0000
1.500 0.0894 0.01303 0.00553 0.0063 0.9902 1.0000
1.750 0.1395 0.01341 0.00605 0.0009 0.9794 1.0000
2.000 0.1923 0.01374 0.00659 -0.0049 0.9681 1.0000
2.250 0.2474 0.01397 0.00706 -0.0109 0.9550 1.0000
2.500 0.3452 0.01294 0.00658 -0.0212 0.9077 1.0000
2.750 0.3726 0.01215 0.00585 -0.0167 0.8366 1.0000
3.000 0.3875 0.01194 0.00558 -0.0115 0.7642 1.0000
3.250 0.3864 0.01327 0.00499 -0.0030 0.3233 1.0000
3.500 0.3977 0.01620 0.00644 -0.0007 0.1023 1.0000
3.750 0.4172 0.01748 0.00763 0.0009 0.0853 1.0000
4.000 0.4388 0.01873 0.00888 0.0023 0.0773 1.0000
4.250 0.4623 0.02054 0.01058 0.0035 0.0728 1.0000
4.500 0.4872 0.02173 0.01191 0.0044 0.0659 1.0000
4.750 0.5128 0.02356 0.01384 0.0052 0.0634 1.0000
5.000 0.5389 0.02571 0.01623 0.0062 0.0633 1.0000
5.250 0.5647 0.02830 0.01939 0.0077 0.0664 1.0000
5.500 0.5886 0.03161 0.02314 0.0090 0.0710 1.0000
5.750 0.6102 0.03502 0.02693 0.0101 0.0725 1.0000
7.500 0.7033 0.06840 0.06365 0.0107 0.1350 1.0000
7.750 0.7296 0.07379 0.06868 0.0137 0.1276 1.0000
8.000 0.7025 0.07717 0.07257 0.0074 0.1205 1.0000
8.250 0.7272 0.08167 0.07687 0.0116 0.1137 1.0000
8.500 0.6996 0.08639 0.08181 0.0043 0.1119 1.0000
8.750 0.6830 0.09248 0.08785 -0.0039 0.1091 1.0000
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Polar data table (+)
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