E216 (10.4%) (e216-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: E216 (10.4%) (e216-il) Reynolds number: 500,000 Max Cl/Cd: 115.85 at α=2° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e216-il-500000-n5.txt Download as CSV file: xf-e216-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: E216 (10.4%)
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.2163 0.08410 0.08163 -0.0823 0.9760 0.0080
-8.750 -0.2197 0.07989 0.07745 -0.0834 0.9701 0.0079
-8.250 -0.2952 0.03183 0.02881 -0.1372 0.9438 0.0073
-8.000 -0.2576 0.02515 0.02141 -0.1495 0.9421 0.0075
-7.750 -0.2302 0.02225 0.01809 -0.1530 0.9382 0.0077
-7.500 -0.1995 0.02022 0.01571 -0.1557 0.9351 0.0079
-7.250 -0.1659 0.01868 0.01388 -0.1582 0.9333 0.0081
-7.000 -0.1316 0.01740 0.01237 -0.1606 0.9319 0.0082
-6.750 -0.0966 0.01611 0.01089 -0.1630 0.9308 0.0085
-6.500 -0.0617 0.01507 0.00976 -0.1653 0.9297 0.0090
-6.250 -0.0263 0.01425 0.00883 -0.1674 0.9288 0.0093
-6.000 0.0046 0.01352 0.00801 -0.1685 0.9258 0.0096
-5.750 0.0341 0.01288 0.00729 -0.1692 0.9216 0.0099
-5.500 0.0684 0.01225 0.00657 -0.1708 0.9193 0.0103
-5.250 0.1040 0.01165 0.00588 -0.1727 0.9173 0.0107
-5.000 0.1406 0.01114 0.00529 -0.1747 0.9156 0.0112
-4.750 0.1792 0.01057 0.00465 -0.1773 0.9143 0.0119
-4.500 0.2072 0.01026 0.00429 -0.1774 0.9085 0.0128
-4.250 0.2415 0.00993 0.00390 -0.1788 0.9048 0.0143
-4.000 0.2798 0.00954 0.00348 -0.1812 0.9021 0.0173
-3.750 0.3207 0.00911 0.00308 -0.1841 0.8997 0.0310
-3.500 0.3513 0.00883 0.00287 -0.1848 0.8930 0.0552
-3.250 0.3905 0.00848 0.00263 -0.1875 0.8884 0.0907
-3.000 0.4320 0.00816 0.00243 -0.1906 0.8841 0.1347
-2.750 0.4660 0.00795 0.00231 -0.1921 0.8764 0.1731
-2.250 0.5389 0.00761 0.00208 -0.1960 0.8606 0.2453
-2.000 0.5735 0.00750 0.00200 -0.1975 0.8513 0.2799
-1.750 0.6073 0.00742 0.00193 -0.1988 0.8408 0.3088
-1.500 0.6378 0.00739 0.00190 -0.1993 0.8287 0.3336
-1.250 0.6681 0.00737 0.00188 -0.1998 0.8164 0.3593
-1.000 0.6978 0.00737 0.00188 -0.2002 0.8034 0.3847
-0.750 0.7266 0.00739 0.00189 -0.2004 0.7895 0.4105
-0.500 0.7547 0.00744 0.00191 -0.2004 0.7748 0.4347
-0.250 0.7820 0.00750 0.00195 -0.2002 0.7590 0.4575
0.000 0.8088 0.00758 0.00201 -0.1999 0.7425 0.4800
0.250 0.8350 0.00768 0.00208 -0.1995 0.7253 0.5025
0.500 0.8609 0.00779 0.00216 -0.1990 0.7078 0.5245
0.750 0.8863 0.00792 0.00225 -0.1984 0.6900 0.5466
1.250 0.9366 0.00821 0.00249 -0.1972 0.6541 0.5907
1.500 0.9615 0.00836 0.00262 -0.1965 0.6358 0.6131
1.750 0.9861 0.00853 0.00276 -0.1958 0.6169 0.6349
2.000 1.0102 0.00872 0.00293 -0.1950 0.5972 0.6568
2.250 1.0340 0.00893 0.00310 -0.1941 0.5766 0.6788
2.500 1.0579 0.00914 0.00328 -0.1933 0.5556 0.7004
2.750 1.0814 0.00937 0.00347 -0.1924 0.5352 0.7226
3.000 1.1047 0.00960 0.00369 -0.1914 0.5153 0.7451
3.250 1.1280 0.00983 0.00390 -0.1905 0.4952 0.7681
3.500 1.1505 0.01008 0.00412 -0.1894 0.4742 0.7919
3.750 1.1721 0.01035 0.00437 -0.1881 0.4526 0.8165
4.000 1.1931 0.01063 0.00462 -0.1866 0.4294 0.8424
4.500 1.2307 0.01113 0.00508 -0.1828 0.3889 0.9077
4.750 1.2468 0.01132 0.00525 -0.1803 0.3687 1.0000
5.000 1.2689 0.01172 0.00555 -0.1793 0.3471 1.0000
5.250 1.2911 0.01211 0.00586 -0.1783 0.3256 1.0000
5.500 1.3124 0.01256 0.00621 -0.1771 0.3024 1.0000
5.750 1.3328 0.01307 0.00659 -0.1758 0.2755 1.0000
6.000 1.3511 0.01373 0.00706 -0.1742 0.2411 1.0000
6.250 1.3694 0.01440 0.00754 -0.1726 0.2084 1.0000
6.500 1.3881 0.01502 0.00803 -0.1711 0.1824 1.0000
6.750 1.4074 0.01559 0.00851 -0.1696 0.1618 1.0000
7.000 1.4251 0.01626 0.00906 -0.1679 0.1390 1.0000
7.250 1.4429 0.01690 0.00961 -0.1663 0.1193 1.0000
7.500 1.4604 0.01756 0.01020 -0.1646 0.1022 1.0000
7.750 1.4769 0.01829 0.01084 -0.1627 0.0855 1.0000
8.000 1.4927 0.01907 0.01153 -0.1608 0.0681 1.0000
8.250 1.5077 0.01988 0.01227 -0.1588 0.0538 1.0000
8.500 1.5227 0.02070 0.01303 -0.1567 0.0426 1.0000
8.750 1.5374 0.02153 0.01384 -0.1547 0.0329 1.0000
9.000 1.5508 0.02245 0.01474 -0.1525 0.0234 1.0000
9.250 1.5628 0.02350 0.01575 -0.1501 0.0150 1.0000
9.500 1.5756 0.02447 0.01674 -0.1479 0.0121 1.0000
9.750 1.5885 0.02544 0.01776 -0.1458 0.0106 1.0000
10.000 1.6017 0.02638 0.01878 -0.1437 0.0099 1.0000
10.250 1.6145 0.02737 0.01985 -0.1417 0.0093 1.0000
10.500 1.6261 0.02846 0.02102 -0.1395 0.0087 1.0000
10.750 1.6365 0.02967 0.02231 -0.1373 0.0083 1.0000
11.000 1.6456 0.03101 0.02374 -0.1351 0.0079 1.0000
11.250 1.6540 0.03243 0.02525 -0.1328 0.0077 1.0000
11.500 1.6630 0.03383 0.02677 -0.1307 0.0076 1.0000
11.750 1.6708 0.03535 0.02840 -0.1286 0.0074 1.0000
12.000 1.6777 0.03700 0.03015 -0.1266 0.0072 1.0000
12.250 1.6835 0.03880 0.03206 -0.1246 0.0071 1.0000
12.500 1.6881 0.04076 0.03413 -0.1226 0.0069 1.0000
12.750 1.6917 0.04288 0.03637 -0.1208 0.0068 1.0000
13.000 1.6943 0.04517 0.03878 -0.1190 0.0067 1.0000
13.250 1.6958 0.04768 0.04141 -0.1174 0.0066 1.0000
13.500 1.6967 0.05037 0.04422 -0.1160 0.0064 1.0000
13.750 1.6963 0.05331 0.04728 -0.1148 0.0063 1.0000
14.000 1.6949 0.05648 0.05059 -0.1138 0.0062 1.0000
14.250 1.6927 0.05991 0.05414 -0.1131 0.0061 1.0000
14.500 1.6890 0.06367 0.05802 -0.1126 0.0060 1.0000
14.750 1.6838 0.06780 0.06229 -0.1124 0.0059 1.0000
15.000 1.6765 0.07241 0.06702 -0.1126 0.0058 1.0000
15.250 1.6680 0.07740 0.07216 -0.1131 0.0058 1.0000
15.500 1.6567 0.08302 0.07792 -0.1141 0.0056 1.0000
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Polar data table (+)
Polar graphs
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