NACA 64-008A AIRFOIL (n64008a-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 64-008A AIRFOIL (n64008a-il) Reynolds number: 1,000,000 Max Cl/Cd: 49.86 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n64008a-il-1000000.txt Download as CSV file: xf-n64008a-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.6902 0.11121 0.10964 0.0200 1.0000 0.0096
-10.250 -0.6903 0.10609 0.10454 0.0175 1.0000 0.0097
-10.000 -0.6912 0.10073 0.09919 0.0146 1.0000 0.0097
-9.750 -0.6929 0.09508 0.09355 0.0114 1.0000 0.0097
-9.500 -0.6956 0.08891 0.08740 0.0073 1.0000 0.0097
-9.250 -0.6995 0.08107 0.07958 0.0002 1.0000 0.0097
-7.000 -0.7245 0.02430 0.02045 -0.0074 1.0000 0.0078
-6.750 -0.7092 0.01968 0.01530 -0.0052 1.0000 0.0080
-6.500 -0.6872 0.01797 0.01334 -0.0042 1.0000 0.0086
-6.250 -0.6641 0.01677 0.01196 -0.0034 1.0000 0.0092
-6.000 -0.6404 0.01590 0.01097 -0.0026 1.0000 0.0095
-5.750 -0.6154 0.01578 0.01079 -0.0021 1.0000 0.0098
-5.500 -0.5988 0.01245 0.00716 0.0000 1.0000 0.0107
-5.250 -0.5762 0.01184 0.00653 0.0008 1.0000 0.0118
-5.000 -0.5532 0.01135 0.00601 0.0016 1.0000 0.0127
-4.750 -0.5307 0.01078 0.00539 0.0027 1.0000 0.0135
-4.500 -0.5086 0.01024 0.00480 0.0037 1.0000 0.0141
-4.250 -0.4865 0.00976 0.00426 0.0048 1.0000 0.0147
-4.000 -0.4533 0.00934 0.00379 0.0034 0.9978 0.0154
-3.750 -0.4194 0.00860 0.00295 0.0019 0.9949 0.0174
-3.500 -0.3858 0.00819 0.00249 0.0005 0.9912 0.0201
-3.250 -0.3516 0.00791 0.00219 -0.0010 0.9871 0.0230
-3.000 -0.3169 0.00755 0.00186 -0.0026 0.9829 0.0385
-2.750 -0.2871 0.00695 0.00160 -0.0034 0.9743 0.1215
-2.500 -0.2592 0.00595 0.00128 -0.0042 0.9646 0.3090
-2.250 -0.2339 0.00510 0.00107 -0.0042 0.9519 0.4923
-2.000 -0.2078 0.00480 0.00097 -0.0039 0.9375 0.5641
-1.750 -0.1823 0.00459 0.00091 -0.0033 0.9220 0.6202
-1.500 -0.1567 0.00445 0.00085 -0.0028 0.9065 0.6629
-1.250 -0.1314 0.00433 0.00083 -0.0021 0.8910 0.7061
-1.000 -0.1060 0.00424 0.00082 -0.0015 0.8760 0.7454
-0.750 -0.0799 0.00420 0.00080 -0.0010 0.8617 0.7696
-0.500 -0.0534 0.00418 0.00079 -0.0006 0.8478 0.7897
-0.250 -0.0268 0.00416 0.00079 -0.0003 0.8343 0.8069
0.000 0.0000 0.00415 0.00078 0.0000 0.8212 0.8212
0.250 0.0269 0.00416 0.00079 0.0003 0.8069 0.8342
0.500 0.0534 0.00418 0.00079 0.0006 0.7898 0.8478
0.750 0.0799 0.00420 0.00080 0.0010 0.7697 0.8617
1.000 0.1060 0.00424 0.00082 0.0015 0.7456 0.8760
1.250 0.1314 0.00434 0.00083 0.0021 0.7054 0.8910
1.500 0.1568 0.00445 0.00085 0.0027 0.6628 0.9065
1.750 0.1823 0.00459 0.00091 0.0033 0.6203 0.9220
2.000 0.2078 0.00480 0.00097 0.0038 0.5647 0.9374
2.250 0.2339 0.00510 0.00107 0.0042 0.4925 0.9518
2.500 0.2592 0.00595 0.00128 0.0042 0.3104 0.9644
2.750 0.2871 0.00695 0.00160 0.0035 0.1215 0.9742
3.000 0.3169 0.00755 0.00186 0.0027 0.0384 0.9828
3.250 0.3516 0.00791 0.00219 0.0010 0.0230 0.9870
3.500 0.3859 0.00818 0.00248 -0.0005 0.0201 0.9911
3.750 0.4194 0.00860 0.00295 -0.0019 0.0174 0.9948
4.000 0.4533 0.00933 0.00379 -0.0034 0.0154 0.9977
4.250 0.4866 0.00976 0.00427 -0.0048 0.0147 1.0000
4.500 0.5085 0.01025 0.00481 -0.0037 0.0141 1.0000
4.750 0.5307 0.01078 0.00538 -0.0027 0.0135 1.0000
5.000 0.5531 0.01135 0.00602 -0.0016 0.0127 1.0000
5.250 0.5761 0.01184 0.00654 -0.0008 0.0118 1.0000
5.500 0.5987 0.01245 0.00715 0.0001 0.0107 1.0000
5.750 0.6154 0.01571 0.01070 0.0021 0.0098 1.0000
6.000 0.6403 0.01590 0.01097 0.0026 0.0095 1.0000
6.250 0.6641 0.01675 0.01195 0.0033 0.0092 1.0000
6.500 0.6872 0.01797 0.01334 0.0042 0.0086 1.0000
6.750 0.7093 0.01967 0.01528 0.0052 0.0080 1.0000
7.000 0.7245 0.02438 0.02054 0.0074 0.0078 1.0000
7.250 0.7083 0.03798 0.03511 0.0114 0.0098 1.0000
7.500 0.7176 0.04243 0.03985 0.0125 0.0098 1.0000
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Polar data table (+)
Polar graphs
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