E214 (11.1%) (e214-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: E214 (11.1%) (e214-il) Reynolds number: 500,000 Max Cl/Cd: 111.28 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e214-il-500000-n5.txt Download as CSV file: xf-e214-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: E214 (11.1%)
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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.3500 0.09603 0.09371 -0.0530 0.9939 0.0081
-10.000 -0.3452 0.09101 0.08870 -0.0569 0.9916 0.0083
-9.500 -0.4718 0.03703 0.03433 -0.1087 0.9681 0.0083
-9.250 -0.4568 0.03340 0.03046 -0.1132 0.9625 0.0085
-9.000 -0.4338 0.03008 0.02683 -0.1175 0.9598 0.0087
-8.750 -0.4117 0.02950 0.02621 -0.1178 0.9553 0.0091
-8.500 -0.3889 0.02787 0.02437 -0.1192 0.9511 0.0095
-8.250 -0.3640 0.02486 0.02098 -0.1218 0.9483 0.0101
-8.000 -0.3360 0.02215 0.01784 -0.1242 0.9464 0.0108
-7.750 -0.3146 0.02045 0.01584 -0.1242 0.9411 0.0112
-7.500 -0.2868 0.01910 0.01423 -0.1251 0.9376 0.0115
-7.250 -0.2561 0.01754 0.01239 -0.1267 0.9351 0.0119
-7.000 -0.2231 0.01627 0.01099 -0.1287 0.9332 0.0126
-6.750 -0.1933 0.01566 0.01032 -0.1295 0.9294 0.0132
-6.500 -0.1645 0.01499 0.00955 -0.1301 0.9246 0.0139
-6.250 -0.1310 0.01415 0.00857 -0.1317 0.9214 0.0145
-6.000 -0.0958 0.01336 0.00766 -0.1336 0.9187 0.0150
-5.750 -0.0678 0.01274 0.00693 -0.1338 0.9126 0.0154
-5.500 -0.0351 0.01217 0.00627 -0.1351 0.9076 0.0159
-5.250 -0.0013 0.01178 0.00579 -0.1365 0.9029 0.0164
-5.000 0.0272 0.01099 0.00493 -0.1370 0.8954 0.0173
-4.750 0.0603 0.01049 0.00435 -0.1383 0.8893 0.0181
-4.500 0.0891 0.01014 0.00393 -0.1386 0.8805 0.0188
-4.250 0.1198 0.00983 0.00355 -0.1393 0.8723 0.0197
-4.000 0.1498 0.00957 0.00321 -0.1398 0.8629 0.0205
-3.750 0.1786 0.00936 0.00292 -0.1400 0.8522 0.0214
-3.500 0.2076 0.00918 0.00266 -0.1402 0.8411 0.0227
-3.250 0.2363 0.00898 0.00240 -0.1404 0.8290 0.0266
-3.000 0.2647 0.00879 0.00219 -0.1406 0.8163 0.0376
-2.750 0.2925 0.00857 0.00202 -0.1407 0.8029 0.0697
-2.500 0.3200 0.00838 0.00188 -0.1407 0.7893 0.1077
-2.250 0.3473 0.00820 0.00178 -0.1407 0.7755 0.1535
-2.000 0.3744 0.00798 0.00169 -0.1408 0.7620 0.2180
-1.750 0.4014 0.00774 0.00162 -0.1409 0.7489 0.2977
-1.500 0.4284 0.00741 0.00160 -0.1410 0.7361 0.4179
-1.250 0.4553 0.00731 0.00164 -0.1409 0.7240 0.4969
-1.000 0.4820 0.00732 0.00167 -0.1407 0.7121 0.5349
-0.750 0.5088 0.00737 0.00169 -0.1404 0.7008 0.5560
-0.500 0.5355 0.00743 0.00171 -0.1401 0.6892 0.5719
-0.250 0.5623 0.00748 0.00174 -0.1399 0.6780 0.5874
0.000 0.5890 0.00754 0.00178 -0.1396 0.6671 0.6024
0.250 0.6156 0.00761 0.00182 -0.1393 0.6565 0.6167
0.500 0.6420 0.00768 0.00187 -0.1390 0.6455 0.6302
0.750 0.6687 0.00775 0.00193 -0.1387 0.6349 0.6438
1.000 0.6951 0.00782 0.00200 -0.1384 0.6245 0.6576
1.250 0.7214 0.00790 0.00207 -0.1381 0.6141 0.6721
1.500 0.7476 0.00798 0.00215 -0.1378 0.6032 0.6864
1.750 0.7739 0.00806 0.00224 -0.1374 0.5921 0.6996
2.000 0.7998 0.00816 0.00233 -0.1370 0.5806 0.7123
2.500 0.8509 0.00837 0.00254 -0.1361 0.5559 0.7403
3.000 0.9011 0.00859 0.00280 -0.1350 0.5300 0.7743
3.250 0.9257 0.00870 0.00293 -0.1343 0.5165 0.7947
3.500 0.9495 0.00882 0.00308 -0.1335 0.5020 0.8188
3.750 0.9717 0.00895 0.00324 -0.1323 0.4859 0.8500
4.000 0.9907 0.00903 0.00338 -0.1303 0.4696 0.8972
4.250 1.0138 0.00911 0.00348 -0.1293 0.4533 1.0000
4.500 1.0384 0.00935 0.00367 -0.1288 0.4356 1.0000
4.750 1.0624 0.00962 0.00388 -0.1281 0.4167 1.0000
5.000 1.0856 0.00994 0.00411 -0.1273 0.3948 1.0000
5.250 1.1081 0.01030 0.00437 -0.1265 0.3700 1.0000
5.500 1.1309 0.01065 0.00463 -0.1256 0.3474 1.0000
5.750 1.1527 0.01105 0.00493 -0.1247 0.3226 1.0000
6.000 1.1738 0.01151 0.00528 -0.1236 0.2954 1.0000
6.250 1.1940 0.01202 0.00566 -0.1223 0.2657 1.0000
6.500 1.2130 0.01260 0.00608 -0.1209 0.2328 1.0000
6.750 1.2310 0.01323 0.00655 -0.1194 0.2003 1.0000
7.000 1.2498 0.01379 0.00700 -0.1179 0.1759 1.0000
7.250 1.2676 0.01434 0.00747 -0.1163 0.1554 1.0000
7.500 1.2843 0.01492 0.00796 -0.1145 0.1334 1.0000
7.750 1.2994 0.01560 0.00852 -0.1124 0.1107 1.0000
8.000 1.3141 0.01629 0.00910 -0.1103 0.0907 1.0000
8.250 1.3288 0.01697 0.00970 -0.1082 0.0732 1.0000
8.500 1.3432 0.01766 0.01032 -0.1062 0.0593 1.0000
8.750 1.3583 0.01831 0.01095 -0.1042 0.0496 1.0000
9.000 1.3725 0.01900 0.01162 -0.1022 0.0409 1.0000
9.250 1.3864 0.01972 0.01232 -0.1002 0.0338 1.0000
9.500 1.4008 0.02041 0.01303 -0.0983 0.0286 1.0000
9.750 1.4135 0.02121 0.01383 -0.0962 0.0231 1.0000
10.000 1.4248 0.02210 0.01473 -0.0939 0.0180 1.0000
10.250 1.4355 0.02305 0.01570 -0.0917 0.0133 1.0000
10.500 1.4455 0.02408 0.01674 -0.0895 0.0097 1.0000
10.750 1.4551 0.02515 0.01786 -0.0873 0.0075 1.0000
11.000 1.4644 0.02627 0.01905 -0.0852 0.0065 1.0000
11.250 1.4740 0.02741 0.02027 -0.0832 0.0058 1.0000
11.500 1.4828 0.02864 0.02158 -0.0813 0.0054 1.0000
11.750 1.4905 0.03000 0.02303 -0.0793 0.0050 1.0000
12.000 1.4961 0.03158 0.02470 -0.0773 0.0046 1.0000
12.250 1.5015 0.03323 0.02646 -0.0755 0.0044 1.0000
12.500 1.5076 0.03488 0.02823 -0.0739 0.0042 1.0000
12.750 1.5126 0.03669 0.03015 -0.0723 0.0042 1.0000
13.000 1.5164 0.03867 0.03225 -0.0709 0.0040 1.0000
13.250 1.5197 0.04080 0.03450 -0.0696 0.0039 1.0000
13.500 1.5219 0.04311 0.03693 -0.0685 0.0038 1.0000
13.750 1.5237 0.04557 0.03950 -0.0676 0.0036 1.0000
14.000 1.5248 0.04820 0.04224 -0.0669 0.0035 1.0000
14.250 1.5245 0.05110 0.04526 -0.0663 0.0034 1.0000
14.500 1.5237 0.05417 0.04846 -0.0660 0.0033 1.0000
14.750 1.5216 0.05754 0.05195 -0.0659 0.0032 1.0000
15.000 1.5180 0.06124 0.05577 -0.0660 0.0031 1.0000
15.250 1.5124 0.06535 0.06002 -0.0664 0.0030 1.0000
15.500 1.5052 0.06988 0.06468 -0.0672 0.0029 1.0000
15.750 1.4959 0.07490 0.06985 -0.0683 0.0029 1.0000
16.000 1.4867 0.08012 0.07524 -0.0697 0.0028 1.0000
16.250 1.4755 0.08584 0.08110 -0.0715 0.0028 1.0000
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Polar data table (+)
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