E211 (10.96%) (e211-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: E211 (10.96%) (e211-il) Reynolds number: 500,000 Max Cl/Cd: 106.06 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e211-il-500000.txt Download as CSV file: xf-e211-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: E211 (10.96%)
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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4713 0.08633 0.08415 -0.0359 1.0000 0.0250
-9.000 -0.4787 0.08231 0.08016 -0.0368 1.0000 0.0256
-8.750 -0.4909 0.07767 0.07557 -0.0381 1.0000 0.0256
-8.500 -0.5087 0.07271 0.07066 -0.0395 1.0000 0.0260
-8.250 -0.5502 0.06533 0.06333 -0.0432 1.0000 0.0251
-8.000 -0.5734 0.05892 0.05681 -0.0468 0.9998 0.0251
-7.750 -0.4909 0.03660 0.03388 -0.0642 0.9880 0.0293
-7.500 -0.4906 0.02550 0.02249 -0.0697 0.9850 0.0309
-7.250 -0.4689 0.02312 0.02006 -0.0712 0.9813 0.0319
-7.000 -0.4955 0.02659 0.02216 -0.0713 0.9837 0.0208
-6.750 -0.4722 0.02137 0.01621 -0.0717 0.9798 0.0183
-6.500 -0.4403 0.01973 0.01435 -0.0731 0.9772 0.0186
-6.250 -0.4072 0.01776 0.01211 -0.0744 0.9754 0.0180
-6.000 -0.3724 0.01645 0.01063 -0.0761 0.9739 0.0179
-5.750 -0.3363 0.01538 0.00948 -0.0781 0.9726 0.0185
-5.500 -0.3094 0.01453 0.00857 -0.0781 0.9676 0.0189
-5.250 -0.2761 0.01376 0.00774 -0.0794 0.9645 0.0195
-5.000 -0.2403 0.01307 0.00698 -0.0813 0.9622 0.0203
-4.750 -0.2030 0.01247 0.00631 -0.0834 0.9605 0.0213
-4.500 -0.1706 0.01195 0.00574 -0.0844 0.9568 0.0226
-4.250 -0.1399 0.01133 0.00508 -0.0851 0.9517 0.0256
-4.000 -0.1035 0.01083 0.00456 -0.0869 0.9491 0.0316
-3.750 -0.0661 0.01000 0.00405 -0.0893 0.9472 0.0978
-3.500 -0.0373 0.00945 0.00372 -0.0897 0.9411 0.1589
-3.250 -0.0035 0.00880 0.00340 -0.0913 0.9370 0.2443
-3.000 0.0326 0.00802 0.00308 -0.0935 0.9344 0.3681
-2.750 0.0623 0.00751 0.00287 -0.0940 0.9282 0.4649
-2.500 0.0954 0.00719 0.00273 -0.0950 0.9226 0.5337
-2.250 0.1347 0.00698 0.00256 -0.0973 0.9191 0.5706
-2.000 0.1659 0.00684 0.00243 -0.0978 0.9091 0.5938
-1.750 0.2039 0.00670 0.00229 -0.0998 0.9011 0.6127
-1.500 0.2435 0.00657 0.00214 -0.1022 0.8916 0.6269
-1.250 0.2799 0.00650 0.00202 -0.1038 0.8778 0.6399
-1.000 0.3145 0.00647 0.00193 -0.1051 0.8613 0.6521
-0.750 0.3455 0.00649 0.00188 -0.1056 0.8421 0.6632
-0.500 0.3749 0.00653 0.00185 -0.1058 0.8222 0.6730
-0.250 0.4031 0.00661 0.00184 -0.1057 0.8027 0.6831
0.000 0.4301 0.00669 0.00186 -0.1054 0.7831 0.6933
0.250 0.4565 0.00677 0.00189 -0.1049 0.7644 0.7028
0.500 0.4829 0.00687 0.00192 -0.1045 0.7468 0.7128
0.750 0.5091 0.00697 0.00196 -0.1041 0.7300 0.7232
1.000 0.5350 0.00705 0.00202 -0.1036 0.7136 0.7329
1.250 0.5610 0.00714 0.00208 -0.1031 0.6973 0.7431
1.500 0.5869 0.00724 0.00214 -0.1026 0.6809 0.7539
1.750 0.6125 0.00732 0.00222 -0.1021 0.6649 0.7638
2.000 0.6381 0.00741 0.00230 -0.1016 0.6494 0.7747
2.250 0.6638 0.00750 0.00239 -0.1011 0.6345 0.7866
2.500 0.6892 0.00760 0.00248 -0.1005 0.6194 0.7997
2.750 0.7143 0.00769 0.00258 -0.0998 0.6041 0.8140
3.000 0.7389 0.00777 0.00269 -0.0991 0.5888 0.8297
3.250 0.7629 0.00785 0.00281 -0.0982 0.5733 0.8481
3.500 0.7861 0.00793 0.00292 -0.0971 0.5578 0.8706
3.750 0.8073 0.00798 0.00303 -0.0955 0.5423 0.8998
4.000 0.8278 0.00798 0.00311 -0.0937 0.5266 0.9505
4.500 0.8845 0.00834 0.00337 -0.0942 0.4867 1.0000
4.750 0.9090 0.00860 0.00354 -0.0937 0.4599 1.0000
5.000 0.9328 0.00890 0.00374 -0.0930 0.4286 1.0000
5.250 0.9552 0.00931 0.00397 -0.0921 0.3879 1.0000
5.500 0.9768 0.00980 0.00425 -0.0910 0.3408 1.0000
5.750 0.9970 0.01044 0.00462 -0.0898 0.2855 1.0000
6.250 1.0366 0.01185 0.00552 -0.0874 0.1853 1.0000
6.500 1.0562 0.01256 0.00601 -0.0862 0.1433 1.0000
6.750 1.0755 0.01329 0.00652 -0.0849 0.1069 1.0000
7.000 1.0953 0.01395 0.00705 -0.0837 0.0794 1.0000
7.250 1.1134 0.01475 0.00766 -0.0822 0.0484 1.0000
7.500 1.1275 0.01588 0.00855 -0.0800 0.0204 1.0000
7.750 1.1467 0.01655 0.00924 -0.0785 0.0165 1.0000
8.000 1.1643 0.01734 0.01009 -0.0767 0.0147 1.0000
8.250 1.1826 0.01800 0.01085 -0.0752 0.0138 1.0000
8.500 1.1992 0.01872 0.01165 -0.0733 0.0131 1.0000
8.750 1.2139 0.01946 0.01245 -0.0711 0.0124 1.0000
9.000 1.2264 0.02032 0.01336 -0.0687 0.0119 1.0000
9.250 1.2345 0.02146 0.01458 -0.0655 0.0113 1.0000
9.500 1.2385 0.02291 0.01614 -0.0619 0.0110 1.0000
9.750 1.2497 0.02390 0.01722 -0.0595 0.0108 1.0000
10.000 1.2591 0.02505 0.01847 -0.0570 0.0107 1.0000
10.250 1.2689 0.02620 0.01973 -0.0546 0.0105 1.0000
10.500 1.2777 0.02750 0.02114 -0.0522 0.0104 1.0000
10.750 1.2859 0.02893 0.02267 -0.0499 0.0102 1.0000
11.000 1.2938 0.03045 0.02431 -0.0477 0.0101 1.0000
11.250 1.3013 0.03211 0.02610 -0.0456 0.0099 1.0000
11.500 1.3084 0.03390 0.02802 -0.0436 0.0098 1.0000
11.750 1.3149 0.03580 0.03006 -0.0418 0.0097 1.0000
12.000 1.3205 0.03788 0.03229 -0.0400 0.0096 1.0000
12.250 1.3246 0.04021 0.03479 -0.0383 0.0096 1.0000
12.500 1.3268 0.04282 0.03761 -0.0366 0.0095 1.0000
12.750 1.3265 0.04570 0.04069 -0.0351 0.0095 1.0000
13.000 1.3229 0.04900 0.04422 -0.0338 0.0096 1.0000
13.250 1.3162 0.05275 0.04821 -0.0328 0.0097 1.0000
13.500 1.3060 0.05699 0.05269 -0.0322 0.0098 1.0000
13.750 1.2927 0.06178 0.05773 -0.0323 0.0099 1.0000
14.000 1.2775 0.06702 0.06320 -0.0330 0.0100 1.0000
14.250 1.2606 0.07275 0.06915 -0.0344 0.0102 1.0000
14.500 1.2415 0.07918 0.07579 -0.0367 0.0103 1.0000
14.750 1.2220 0.08609 0.08290 -0.0397 0.0104 1.0000
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Polar data table (+)
Polar graphs
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