EPPLER 678 AIRFOIL (e678-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 678 AIRFOIL (e678-il) Reynolds number: 500,000 Max Cl/Cd: 116.76 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e678-il-500000.txt Download as CSV file: xf-e678-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 678 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.000 0.0458 0.10761 0.10504 -0.1371 0.9519 0.0205
-12.750 0.0558 0.10455 0.10198 -0.1391 0.9515 0.0214
-12.500 -0.0577 0.11392 0.11124 -0.1332 0.9570 0.0199
-12.250 -0.0484 0.11146 0.10878 -0.1340 0.9554 0.0202
-12.000 -0.0371 0.10898 0.10630 -0.1353 0.9542 0.0207
-11.750 -0.0277 0.10585 0.10317 -0.1373 0.9532 0.0212
-11.500 -0.0182 0.10226 0.09958 -0.1401 0.9524 0.0225
-11.000 0.0794 0.07911 0.07656 -0.1495 0.9434 0.0255
-10.750 0.0855 0.07591 0.07337 -0.1507 0.9419 0.0260
-10.500 0.0916 0.07202 0.06948 -0.1529 0.9409 0.0268
-10.250 0.0949 0.06744 0.06489 -0.1555 0.9398 0.0279
-10.000 0.0793 0.05819 0.05567 -0.1623 0.9386 0.0296
-9.750 0.0183 0.07562 0.07299 -0.1571 0.9410 0.0265
-9.500 0.0263 0.07100 0.06838 -0.1616 0.9396 0.0273
-9.250 0.0302 0.06393 0.06131 -0.1694 0.9381 0.0285
-7.750 -0.1252 0.02917 0.02471 -0.1680 0.8993 0.0196
-7.500 -0.1012 0.02452 0.01981 -0.1701 0.8975 0.0180
-7.250 -0.0702 0.02094 0.01570 -0.1723 0.8961 0.0172
-7.000 -0.0337 0.01900 0.01348 -0.1748 0.8949 0.0173
-6.750 0.0050 0.01756 0.01184 -0.1776 0.8938 0.0177
-6.500 0.0144 0.01692 0.01111 -0.1741 0.8879 0.0179
-6.250 0.0407 0.01611 0.01021 -0.1741 0.8842 0.0185
-6.000 0.0766 0.01531 0.00929 -0.1760 0.8819 0.0193
-5.750 0.1141 0.01423 0.00812 -0.1785 0.8798 0.0198
-5.500 0.1424 0.01331 0.00717 -0.1791 0.8761 0.0206
-5.250 0.1592 0.01283 0.00669 -0.1771 0.8693 0.0215
-5.000 0.1951 0.01221 0.00600 -0.1791 0.8656 0.0232
-4.750 0.2237 0.01163 0.00537 -0.1796 0.8604 0.0251
-4.500 0.2452 0.01122 0.00494 -0.1785 0.8524 0.0281
-4.250 0.2806 0.01065 0.00437 -0.1804 0.8474 0.0398
-4.000 0.2965 0.01014 0.00402 -0.1782 0.8362 0.0752
-3.750 0.3160 0.00975 0.00376 -0.1768 0.8238 0.1139
-3.500 0.3369 0.00935 0.00350 -0.1757 0.8087 0.1640
-3.250 0.3663 0.00882 0.00320 -0.1767 0.7922 0.2521
-3.000 0.4019 0.00817 0.00293 -0.1792 0.7764 0.3994
-2.750 0.4359 0.00780 0.00280 -0.1811 0.7630 0.5184
-2.500 0.4666 0.00771 0.00278 -0.1819 0.7530 0.5854
-2.250 0.4966 0.00777 0.00283 -0.1825 0.7457 0.6298
-2.000 0.5259 0.00789 0.00292 -0.1829 0.7398 0.6616
-1.750 0.5568 0.00807 0.00301 -0.1836 0.7350 0.6875
-1.500 0.5856 0.00822 0.00312 -0.1839 0.7309 0.7040
-1.250 0.6135 0.00837 0.00324 -0.1840 0.7269 0.7159
-1.000 0.6431 0.00854 0.00335 -0.1845 0.7230 0.7272
-0.750 0.6751 0.00874 0.00346 -0.1855 0.7195 0.7376
-0.500 0.7037 0.00890 0.00361 -0.1858 0.7166 0.7459
-0.250 0.7312 0.00905 0.00373 -0.1859 0.7137 0.7556
0.000 0.7577 0.00918 0.00388 -0.1856 0.7107 0.7621
0.250 0.7868 0.00933 0.00399 -0.1861 0.7078 0.7694
0.500 0.8168 0.00948 0.00410 -0.1867 0.7050 0.7745
0.750 0.8492 0.00968 0.00426 -0.1879 0.7023 0.7797
1.000 0.8757 0.00979 0.00438 -0.1879 0.6998 0.7857
1.250 0.9007 0.00988 0.00450 -0.1874 0.6967 0.7903
1.500 0.9267 0.01000 0.00463 -0.1872 0.6930 0.7950
1.750 0.9555 0.01014 0.00474 -0.1876 0.6891 0.8000
2.000 0.9877 0.01034 0.00489 -0.1888 0.6855 0.8039
2.250 1.0094 0.01039 0.00501 -0.1877 0.6822 0.8075
2.500 1.0341 0.01048 0.00512 -0.1872 0.6786 0.8115
2.750 1.0615 0.01059 0.00523 -0.1874 0.6749 0.8156
3.000 1.0913 0.01073 0.00534 -0.1881 0.6714 0.8191
3.250 1.1165 0.01084 0.00549 -0.1878 0.6677 0.8221
3.500 1.1382 0.01090 0.00562 -0.1867 0.6638 0.8258
3.750 1.1631 0.01098 0.00572 -0.1863 0.6598 0.8296
4.000 1.1911 0.01109 0.00582 -0.1867 0.6560 0.8333
4.250 1.2163 0.01120 0.00596 -0.1864 0.6520 0.8364
4.500 1.2344 0.01123 0.00609 -0.1845 0.6472 0.8401
4.750 1.2564 0.01128 0.00617 -0.1835 0.6423 0.8443
5.000 1.2848 0.01141 0.00628 -0.1839 0.6378 0.8481
5.250 1.2984 0.01142 0.00640 -0.1811 0.6328 0.8518
5.500 1.3143 0.01145 0.00649 -0.1788 0.6275 0.8558
5.750 1.3378 0.01155 0.00658 -0.1781 0.6223 0.8595
6.000 1.3508 0.01165 0.00678 -0.1753 0.6161 0.8639
6.250 1.3658 0.01172 0.00691 -0.1728 0.6096 0.8676
6.500 1.3824 0.01184 0.00709 -0.1708 0.6029 0.8717
6.750 1.3964 0.01198 0.00729 -0.1682 0.5946 0.8765
7.000 1.4098 0.01214 0.00751 -0.1656 0.5849 0.8811
7.250 1.4209 0.01235 0.00776 -0.1625 0.5730 0.8860
7.500 1.4304 0.01264 0.00808 -0.1592 0.5566 0.8917
7.750 1.4352 0.01306 0.00847 -0.1550 0.5325 0.8979
8.000 1.4375 0.01371 0.00902 -0.1506 0.5010 0.9057
8.250 1.4372 0.01455 0.00973 -0.1458 0.4658 0.9142
8.500 1.4355 0.01558 0.01061 -0.1410 0.4301 0.9253
8.750 1.4336 0.01663 0.01154 -0.1363 0.3967 0.9460
9.000 1.4349 0.01784 0.01260 -0.1325 0.3622 1.0000
9.250 1.4403 0.01914 0.01375 -0.1297 0.3317 1.0000
9.500 1.4468 0.02041 0.01491 -0.1272 0.3036 1.0000
9.750 1.4519 0.02180 0.01616 -0.1245 0.2755 1.0000
10.000 1.4558 0.02329 0.01753 -0.1217 0.2483 1.0000
10.250 1.4586 0.02490 0.01900 -0.1189 0.2202 1.0000
10.500 1.4617 0.02656 0.02053 -0.1163 0.1958 1.0000
10.750 1.4654 0.02824 0.02210 -0.1138 0.1718 1.0000
11.000 1.4689 0.02999 0.02375 -0.1115 0.1500 1.0000
11.250 1.4717 0.03186 0.02552 -0.1092 0.1287 1.0000
11.500 1.4747 0.03379 0.02737 -0.1071 0.1091 1.0000
11.750 1.4771 0.03586 0.02934 -0.1051 0.0898 1.0000
12.000 1.4777 0.03815 0.03153 -0.1031 0.0698 1.0000
12.250 1.4758 0.04079 0.03403 -0.1010 0.0486 1.0000
12.500 1.4705 0.04388 0.03697 -0.0989 0.0294 1.0000
12.750 1.4716 0.04645 0.03952 -0.0974 0.0214 1.0000
13.000 1.4747 0.04888 0.04198 -0.0961 0.0185 1.0000
13.250 1.4815 0.05098 0.04416 -0.0952 0.0171 1.0000
13.500 1.4866 0.05328 0.04654 -0.0943 0.0160 1.0000
13.750 1.4885 0.05598 0.04930 -0.0933 0.0149 1.0000
14.000 1.4923 0.05851 0.05192 -0.0925 0.0143 1.0000
14.250 1.4968 0.06099 0.05450 -0.0918 0.0138 1.0000
14.500 1.5004 0.06364 0.05724 -0.0911 0.0133 1.0000
14.750 1.5025 0.06649 0.06018 -0.0906 0.0130 1.0000
15.000 1.5040 0.06948 0.06327 -0.0901 0.0126 1.0000
15.250 1.5041 0.07269 0.06656 -0.0897 0.0123 1.0000
15.500 1.5020 0.07624 0.07022 -0.0893 0.0120 1.0000
15.750 1.4972 0.08025 0.07433 -0.0891 0.0118 1.0000
16.000 1.4902 0.08463 0.07882 -0.0890 0.0115 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 678 AIRFOIL (e678-il)