E212 (10.55%) (e212-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: E212 (10.55%) (e212-il) Reynolds number: 500,000 Max Cl/Cd: 108.43 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e212-il-500000.txt Download as CSV file: xf-e212-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: E212 (10.55%)
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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4612 0.08966 0.08743 -0.0359 1.0000 0.0281
-9.250 -0.4687 0.08544 0.08325 -0.0369 1.0000 0.0289
-8.250 -0.6134 0.03452 0.03099 -0.0724 0.9930 0.0188
-8.000 -0.5936 0.02467 0.01995 -0.0776 0.9895 0.0157
-7.750 -0.5619 0.02246 0.01742 -0.0798 0.9876 0.0156
-7.500 -0.5287 0.02068 0.01540 -0.0818 0.9862 0.0157
-7.250 -0.4946 0.01913 0.01369 -0.0839 0.9850 0.0159
-7.000 -0.4643 0.01799 0.01246 -0.0850 0.9822 0.0163
-6.750 -0.4327 0.01705 0.01143 -0.0863 0.9794 0.0169
-6.500 -0.3987 0.01601 0.01028 -0.0880 0.9772 0.0172
-6.250 -0.3632 0.01507 0.00925 -0.0899 0.9755 0.0177
-6.000 -0.3265 0.01428 0.00837 -0.0921 0.9741 0.0183
-5.750 -0.2890 0.01360 0.00761 -0.0944 0.9730 0.0190
-5.500 -0.2607 0.01292 0.00686 -0.0947 0.9682 0.0198
-5.250 -0.2258 0.01223 0.00610 -0.0964 0.9655 0.0217
-5.000 -0.1894 0.01179 0.00561 -0.0983 0.9635 0.0243
-4.750 -0.1517 0.01117 0.00503 -0.1005 0.9619 0.0363
-4.500 -0.1135 0.01034 0.00454 -0.1032 0.9608 0.1056
-4.250 -0.0849 0.00993 0.00426 -0.1035 0.9554 0.1426
-4.000 -0.0500 0.00953 0.00401 -0.1052 0.9523 0.1849
-3.750 -0.0129 0.00914 0.00377 -0.1072 0.9502 0.2286
-3.500 0.0253 0.00879 0.00354 -0.1095 0.9486 0.2687
-3.250 0.0642 0.00844 0.00333 -0.1119 0.9473 0.3141
-3.000 0.0916 0.00820 0.00317 -0.1117 0.9403 0.3449
-2.750 0.1284 0.00790 0.00295 -0.1136 0.9375 0.3782
-2.500 0.1674 0.00763 0.00275 -0.1158 0.9353 0.4098
-2.250 0.2024 0.00740 0.00259 -0.1172 0.9308 0.4371
-2.000 0.2352 0.00721 0.00244 -0.1181 0.9243 0.4625
-1.750 0.2736 0.00699 0.00227 -0.1202 0.9200 0.4896
-1.500 0.3046 0.00684 0.00218 -0.1207 0.9109 0.5137
-1.250 0.3417 0.00668 0.00205 -0.1225 0.9039 0.5391
-1.000 0.3731 0.00657 0.00199 -0.1231 0.8925 0.5623
-0.750 0.4044 0.00650 0.00193 -0.1236 0.8802 0.5855
-0.500 0.4351 0.00643 0.00189 -0.1240 0.8668 0.6073
-0.250 0.4648 0.00641 0.00187 -0.1242 0.8523 0.6292
0.000 0.4936 0.00640 0.00187 -0.1242 0.8366 0.6502
0.250 0.5217 0.00642 0.00189 -0.1240 0.8201 0.6710
0.500 0.5492 0.00646 0.00191 -0.1237 0.8027 0.6920
0.750 0.5760 0.00651 0.00195 -0.1233 0.7846 0.7122
1.000 0.6022 0.00657 0.00201 -0.1227 0.7651 0.7327
1.250 0.6281 0.00665 0.00208 -0.1221 0.7448 0.7532
1.500 0.6534 0.00675 0.00215 -0.1213 0.7242 0.7729
1.750 0.6783 0.00685 0.00225 -0.1205 0.7024 0.7930
2.000 0.7027 0.00699 0.00234 -0.1196 0.6803 0.8134
2.250 0.7267 0.00710 0.00244 -0.1185 0.6573 0.8332
2.500 0.7497 0.00725 0.00255 -0.1173 0.6343 0.8536
2.750 0.7724 0.00738 0.00267 -0.1160 0.6101 0.8753
3.000 0.7929 0.00751 0.00277 -0.1142 0.5862 0.8986
3.250 0.8113 0.00763 0.00285 -0.1119 0.5616 0.9274
3.500 0.8338 0.00769 0.00288 -0.1106 0.5361 1.0000
3.750 0.8602 0.00796 0.00305 -0.1104 0.5091 1.0000
4.000 0.8861 0.00825 0.00322 -0.1101 0.4821 1.0000
4.250 0.9115 0.00856 0.00342 -0.1097 0.4541 1.0000
4.500 0.9367 0.00889 0.00364 -0.1093 0.4243 1.0000
4.750 0.9610 0.00927 0.00387 -0.1087 0.3891 1.0000
5.000 0.9849 0.00969 0.00413 -0.1080 0.3524 1.0000
5.250 1.0084 0.01016 0.00442 -0.1073 0.3142 1.0000
5.500 1.0319 0.01064 0.00475 -0.1067 0.2788 1.0000
5.750 1.0550 0.01115 0.00509 -0.1060 0.2445 1.0000
6.000 1.0776 0.01172 0.00547 -0.1052 0.2056 1.0000
6.250 1.0999 0.01231 0.00589 -0.1043 0.1726 1.0000
6.500 1.1223 0.01289 0.00632 -0.1035 0.1433 1.0000
6.750 1.1442 0.01350 0.00679 -0.1026 0.1154 1.0000
7.000 1.1659 0.01413 0.00728 -0.1017 0.0913 1.0000
7.250 1.1869 0.01481 0.00783 -0.1007 0.0676 1.0000
7.500 1.2043 0.01584 0.00862 -0.0991 0.0344 1.0000
7.750 1.2224 0.01680 0.00949 -0.0975 0.0198 1.0000
8.000 1.2419 0.01760 0.01030 -0.0961 0.0166 1.0000
8.250 1.2622 0.01826 0.01102 -0.0948 0.0154 1.0000
8.500 1.2817 0.01896 0.01179 -0.0934 0.0145 1.0000
8.750 1.2999 0.01974 0.01265 -0.0919 0.0137 1.0000
9.000 1.3153 0.02070 0.01367 -0.0899 0.0131 1.0000
9.250 1.3252 0.02193 0.01500 -0.0870 0.0126 1.0000
9.500 1.3363 0.02295 0.01612 -0.0844 0.0124 1.0000
9.750 1.3477 0.02394 0.01720 -0.0818 0.0122 1.0000
10.000 1.3579 0.02502 0.01838 -0.0792 0.0119 1.0000
10.250 1.3670 0.02622 0.01968 -0.0765 0.0117 1.0000
10.500 1.3753 0.02751 0.02107 -0.0738 0.0115 1.0000
10.750 1.3834 0.02884 0.02251 -0.0713 0.0113 1.0000
11.000 1.3913 0.03024 0.02401 -0.0689 0.0110 1.0000
11.250 1.3988 0.03169 0.02555 -0.0667 0.0108 1.0000
11.500 1.4054 0.03327 0.02723 -0.0645 0.0105 1.0000
11.750 1.4110 0.03501 0.02908 -0.0624 0.0104 1.0000
12.000 1.4157 0.03695 0.03114 -0.0605 0.0103 1.0000
12.250 1.4197 0.03904 0.03334 -0.0586 0.0102 1.0000
12.500 1.4228 0.04130 0.03574 -0.0570 0.0100 1.0000
12.750 1.4250 0.04377 0.03834 -0.0555 0.0100 1.0000
13.000 1.4261 0.04645 0.04116 -0.0541 0.0099 1.0000
13.250 1.4257 0.04940 0.04427 -0.0530 0.0098 1.0000
13.500 1.4239 0.05260 0.04763 -0.0521 0.0098 1.0000
13.750 1.4201 0.05612 0.05133 -0.0516 0.0098 1.0000
14.000 1.4146 0.05998 0.05536 -0.0513 0.0097 1.0000
14.250 1.4071 0.06421 0.05978 -0.0516 0.0097 1.0000
14.500 1.3981 0.06878 0.06455 -0.0523 0.0097 1.0000
14.750 1.3873 0.07385 0.06980 -0.0535 0.0097 1.0000
15.000 1.3751 0.07937 0.07552 -0.0554 0.0097 1.0000
15.250 1.3615 0.08545 0.08181 -0.0580 0.0098 1.0000
15.500 1.3466 0.09212 0.08868 -0.0612 0.0098 1.0000
15.750 1.3294 0.09962 0.09640 -0.0653 0.0099 1.0000
16.000 1.3093 0.10827 0.10528 -0.0703 0.0100 1.0000
16.250 1.2853 0.11836 0.11562 -0.0766 0.0101 1.0000
16.500 1.2442 0.13366 0.13128 -0.0871 0.0105 1.0000
16.750 1.1954 0.15293 0.15088 -0.1004 0.0109 1.0000
17.000 1.1048 0.19025 0.18844 -0.1238 0.0123 1.0000
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