E216 (10.4%) (e216-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: E216 (10.4%) (e216-il) Reynolds number: 500,000 Max Cl/Cd: 136.84 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e216-il-500000.txt Download as CSV file: xf-e216-il-500000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.2213 0.08376 0.08152 -0.0715 0.9724 0.0242
-7.500 -0.2170 0.07696 0.07473 -0.0796 0.9666 0.0261
-7.250 -0.2138 0.07185 0.06965 -0.0844 0.9601 0.0262
-6.500 -0.1306 0.02377 0.01944 -0.1510 0.9427 0.0166
-6.250 -0.0911 0.02202 0.01737 -0.1544 0.9418 0.0163
-6.000 -0.0507 0.01956 0.01456 -0.1583 0.9413 0.0163
-5.750 -0.0109 0.01735 0.01206 -0.1617 0.9409 0.0165
-5.500 0.0284 0.01576 0.01033 -0.1647 0.9404 0.0169
-5.250 0.0572 0.01503 0.00957 -0.1653 0.9365 0.0179
-5.000 0.0902 0.01435 0.00884 -0.1666 0.9334 0.0188
-4.750 0.1278 0.01356 0.00796 -0.1688 0.9319 0.0195
-4.500 0.1664 0.01285 0.00717 -0.1712 0.9307 0.0205
-4.250 0.2068 0.01204 0.00630 -0.1740 0.9297 0.0225
-4.000 0.2458 0.01150 0.00572 -0.1764 0.9288 0.0257
-3.750 0.2873 0.01067 0.00502 -0.1794 0.9281 0.0573
-3.500 0.3272 0.01007 0.00461 -0.1822 0.9274 0.1122
-3.250 0.3668 0.00960 0.00425 -0.1848 0.9266 0.1544
-3.000 0.3919 0.00937 0.00413 -0.1842 0.9198 0.1949
-2.750 0.4291 0.00895 0.00387 -0.1863 0.9175 0.2478
-2.500 0.4683 0.00860 0.00360 -0.1887 0.9156 0.2905
-2.250 0.5095 0.00826 0.00334 -0.1915 0.9138 0.3284
-2.000 0.5413 0.00806 0.00321 -0.1923 0.9082 0.3625
-1.750 0.5774 0.00784 0.00304 -0.1940 0.9031 0.3979
-1.500 0.6199 0.00761 0.00286 -0.1971 0.8997 0.4320
-1.250 0.6519 0.00749 0.00279 -0.1979 0.8921 0.4611
-1.000 0.6901 0.00733 0.00266 -0.2000 0.8859 0.4917
-0.750 0.7219 0.00724 0.00262 -0.2008 0.8769 0.5198
-0.500 0.7603 0.00714 0.00253 -0.2029 0.8693 0.5486
-0.250 0.7893 0.00711 0.00253 -0.2030 0.8578 0.5744
0.000 0.8199 0.00709 0.00253 -0.2035 0.8462 0.5996
0.250 0.8503 0.00709 0.00254 -0.2039 0.8340 0.6245
0.500 0.8797 0.00711 0.00256 -0.2041 0.8209 0.6495
0.750 0.9078 0.00715 0.00261 -0.2040 0.8068 0.6736
1.000 0.9351 0.00722 0.00267 -0.2037 0.7919 0.6977
1.250 0.9613 0.00730 0.00275 -0.2032 0.7763 0.7209
1.750 1.0117 0.00751 0.00295 -0.2017 0.7427 0.7681
2.000 1.0358 0.00764 0.00306 -0.2007 0.7246 0.7919
2.250 1.0588 0.00779 0.00319 -0.1995 0.7058 0.8154
2.500 1.0810 0.00794 0.00332 -0.1981 0.6867 0.8399
2.750 1.1020 0.00809 0.00346 -0.1965 0.6674 0.8661
3.000 1.1209 0.00823 0.00358 -0.1944 0.6484 0.8953
3.250 1.1344 0.00829 0.00365 -0.1911 0.6297 0.9411
3.500 1.1570 0.00847 0.00375 -0.1900 0.6093 1.0000
3.750 1.1810 0.00875 0.00393 -0.1892 0.5870 1.0000
4.000 1.2042 0.00906 0.00415 -0.1883 0.5647 1.0000
4.250 1.2273 0.00936 0.00436 -0.1873 0.5419 1.0000
4.500 1.2497 0.00970 0.00460 -0.1863 0.5194 1.0000
4.750 1.2722 0.01003 0.00485 -0.1852 0.4961 1.0000
5.000 1.2939 0.01041 0.00514 -0.1840 0.4730 1.0000
5.250 1.3153 0.01079 0.00543 -0.1828 0.4484 1.0000
5.500 1.3359 0.01120 0.00574 -0.1814 0.4237 1.0000
5.750 1.3553 0.01166 0.00608 -0.1798 0.3967 1.0000
6.000 1.3746 0.01215 0.00646 -0.1783 0.3689 1.0000
6.250 1.3931 0.01269 0.00686 -0.1766 0.3388 1.0000
6.500 1.4110 0.01329 0.00730 -0.1748 0.3071 1.0000
6.750 1.4291 0.01389 0.00777 -0.1731 0.2778 1.0000
7.000 1.4474 0.01450 0.00826 -0.1715 0.2505 1.0000
7.250 1.4652 0.01514 0.00878 -0.1697 0.2238 1.0000
7.500 1.4819 0.01584 0.00935 -0.1679 0.1969 1.0000
7.750 1.4974 0.01663 0.00999 -0.1658 0.1686 1.0000
8.000 1.5129 0.01742 0.01066 -0.1638 0.1444 1.0000
8.250 1.5283 0.01822 0.01135 -0.1618 0.1225 1.0000
8.500 1.5432 0.01904 0.01208 -0.1597 0.1040 1.0000
8.750 1.5577 0.01988 0.01285 -0.1576 0.0866 1.0000
9.000 1.5706 0.02085 0.01372 -0.1553 0.0697 1.0000
9.250 1.5823 0.02190 0.01469 -0.1529 0.0521 1.0000
9.750 1.5965 0.02478 0.01735 -0.1467 0.0207 1.0000
10.000 1.6064 0.02601 0.01863 -0.1441 0.0175 1.0000
10.250 1.6175 0.02715 0.01983 -0.1417 0.0160 1.0000
10.500 1.6254 0.02856 0.02131 -0.1391 0.0147 1.0000
10.750 1.6322 0.03008 0.02293 -0.1363 0.0140 1.0000
11.000 1.6409 0.03146 0.02443 -0.1340 0.0135 1.0000
11.250 1.6481 0.03300 0.02608 -0.1316 0.0131 1.0000
11.500 1.6540 0.03468 0.02787 -0.1292 0.0127 1.0000
11.750 1.6588 0.03650 0.02979 -0.1268 0.0123 1.0000
12.000 1.6622 0.03852 0.03192 -0.1245 0.0120 1.0000
12.250 1.6634 0.04082 0.03432 -0.1223 0.0117 1.0000
12.500 1.6622 0.04345 0.03706 -0.1200 0.0114 1.0000
12.750 1.6585 0.04647 0.04020 -0.1179 0.0112 1.0000
13.000 1.6518 0.05001 0.04388 -0.1159 0.0110 1.0000
13.250 1.6474 0.05347 0.04749 -0.1144 0.0108 1.0000
13.500 1.6483 0.05640 0.05056 -0.1133 0.0106 1.0000
13.750 1.6492 0.05944 0.05372 -0.1125 0.0104 1.0000
14.000 1.6485 0.06280 0.05722 -0.1119 0.0103 1.0000
14.250 1.6467 0.06642 0.06098 -0.1114 0.0101 1.0000
14.500 1.6438 0.07032 0.06501 -0.1111 0.0100 1.0000
14.750 1.6411 0.07429 0.06912 -0.1110 0.0099 1.0000
15.000 1.6380 0.07844 0.07340 -0.1112 0.0098 1.0000
15.250 1.6347 0.08270 0.07781 -0.1115 0.0096 1.0000
15.500 1.6306 0.08719 0.08243 -0.1121 0.0095 1.0000
15.750 1.6267 0.09175 0.08713 -0.1129 0.0094 1.0000
16.000 1.6222 0.09648 0.09201 -0.1139 0.0093 1.0000
16.250 1.6174 0.10135 0.09704 -0.1152 0.0093 1.0000
16.500 1.6120 0.10640 0.10224 -0.1167 0.0092 1.0000
16.750 1.6067 0.11151 0.10748 -0.1185 0.0090 1.0000
17.000 1.6004 0.11685 0.11298 -0.1205 0.0090 1.0000
17.250 1.5940 0.12233 0.11861 -0.1228 0.0090 1.0000
17.500 1.5875 0.12790 0.12434 -0.1254 0.0089 1.0000
17.750 1.5797 0.13385 0.13045 -0.1284 0.0089 1.0000
18.000 1.5728 0.13963 0.13635 -0.1316 0.0088 1.0000
18.250 1.5653 0.14566 0.14252 -0.1351 0.0087 1.0000
18.500 1.5548 0.15256 0.14962 -0.1393 0.0088 1.0000
18.750 1.5445 0.15959 0.15683 -0.1439 0.0088 1.0000
19.000 1.5370 0.16600 0.16335 -0.1482 0.0087 1.0000
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Polar data table (+)
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