GOE 611 AIRFOIL (goe611-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 611 AIRFOIL (goe611-il) Reynolds number: 200,000 Max Cl/Cd: 75.77 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe611-il-200000-n5.txt Download as CSV file: xf-goe611-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 611 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.1052 0.08965 0.08616 -0.0847 0.9100 0.0228
-8.500 -0.0980 0.08595 0.08243 -0.0873 0.8971 0.0228
-8.250 -0.0921 0.08230 0.07876 -0.0899 0.8848 0.0228
-8.000 -0.0863 0.07865 0.07507 -0.0930 0.8739 0.0228
-7.750 -0.0770 0.07467 0.07102 -0.0967 0.8641 0.0228
-7.500 -0.0668 0.07071 0.06701 -0.0999 0.8549 0.0228
-7.250 -0.0559 0.06699 0.06320 -0.1021 0.8473 0.0227
-7.000 -0.0475 0.06482 0.06104 -0.1004 0.8395 0.0219
-6.750 -0.0350 0.06099 0.05710 -0.1034 0.8322 0.0212
-6.500 -0.0185 0.05637 0.05236 -0.1081 0.8244 0.0220
-6.250 -0.0011 0.05204 0.04784 -0.1113 0.8177 0.0221
-6.000 0.0158 0.04844 0.04409 -0.1131 0.8107 0.0216
-5.750 0.0349 0.04456 0.04000 -0.1149 0.8040 0.0213
-5.500 0.0558 0.04066 0.03583 -0.1163 0.7981 0.0212
-5.250 0.0781 0.03664 0.03147 -0.1172 0.7911 0.0218
-5.000 0.1014 0.03351 0.02799 -0.1177 0.7853 0.0238
-4.750 0.1243 0.03169 0.02600 -0.1178 0.7781 0.0248
-4.500 0.1489 0.02902 0.02297 -0.1178 0.7717 0.0253
-4.250 0.1744 0.02673 0.02030 -0.1176 0.7651 0.0266
-4.000 0.2022 0.02482 0.01778 -0.1170 0.7580 0.0290
-3.750 0.2276 0.02260 0.01521 -0.1169 0.7517 0.0299
-3.500 0.2533 0.02154 0.01406 -0.1169 0.7437 0.0312
-3.250 0.2801 0.02074 0.01308 -0.1169 0.7367 0.0334
-3.000 0.3073 0.01964 0.01171 -0.1166 0.7284 0.0345
-2.750 0.3350 0.01869 0.01050 -0.1164 0.7210 0.0353
-2.500 0.3625 0.01821 0.00979 -0.1161 0.7122 0.0372
-2.250 0.3901 0.01763 0.00901 -0.1159 0.7040 0.0379
-2.000 0.4170 0.01665 0.00790 -0.1157 0.6952 0.0385
-1.750 0.4437 0.01595 0.00714 -0.1155 0.6861 0.0392
-1.500 0.4705 0.01546 0.00659 -0.1153 0.6774 0.0409
-1.250 0.4972 0.01509 0.00620 -0.1151 0.6674 0.0425
-1.000 0.5239 0.01474 0.00578 -0.1149 0.6580 0.0431
-0.750 0.5505 0.01445 0.00542 -0.1146 0.6483 0.0438
-0.500 0.5772 0.01421 0.00514 -0.1145 0.6383 0.0446
-0.250 0.6039 0.01405 0.00489 -0.1143 0.6285 0.0456
0.000 0.6304 0.01394 0.00468 -0.1141 0.6173 0.0467
0.250 0.6568 0.01390 0.00454 -0.1138 0.6052 0.0488
0.500 0.6831 0.01384 0.00438 -0.1136 0.5937 0.0518
0.750 0.7093 0.01385 0.00429 -0.1133 0.5822 0.0544
1.000 0.7354 0.01389 0.00424 -0.1131 0.5709 0.0586
1.250 0.7617 0.01390 0.00424 -0.1129 0.5605 0.0771
1.750 0.8127 0.01214 0.00449 -0.1121 0.5410 1.0000
2.000 0.8381 0.01233 0.00456 -0.1118 0.5314 1.0000
2.250 0.8632 0.01254 0.00463 -0.1114 0.5226 1.0000
2.500 0.8887 0.01272 0.00475 -0.1111 0.5138 1.0000
2.750 0.9138 0.01294 0.00487 -0.1107 0.5060 1.0000
3.000 0.9391 0.01313 0.00501 -0.1104 0.4975 1.0000
3.250 0.9640 0.01335 0.00516 -0.1100 0.4899 1.0000
3.500 0.9887 0.01356 0.00533 -0.1096 0.4807 1.0000
3.750 1.0131 0.01379 0.00551 -0.1092 0.4717 1.0000
4.000 1.0371 0.01403 0.00570 -0.1087 0.4626 1.0000
4.250 1.0616 0.01426 0.00591 -0.1082 0.4544 1.0000
4.500 1.0853 0.01452 0.00612 -0.1077 0.4464 1.0000
4.750 1.1093 0.01475 0.00636 -0.1072 0.4383 1.0000
5.000 1.1323 0.01503 0.00661 -0.1066 0.4302 1.0000
5.250 1.1558 0.01528 0.00688 -0.1060 0.4216 1.0000
5.500 1.1780 0.01558 0.00715 -0.1053 0.4136 1.0000
5.750 1.2010 0.01585 0.00746 -0.1046 0.4054 1.0000
6.000 1.2226 0.01617 0.00777 -0.1038 0.3978 1.0000
6.250 1.2440 0.01648 0.00810 -0.1029 0.3880 1.0000
6.500 1.2642 0.01683 0.00847 -0.1018 0.3779 1.0000
6.750 1.2833 0.01721 0.00885 -0.1006 0.3688 1.0000
7.000 1.3033 0.01756 0.00924 -0.0995 0.3596 1.0000
7.250 1.3214 0.01797 0.00966 -0.0980 0.3512 1.0000
7.500 1.3380 0.01837 0.01012 -0.0963 0.3421 1.0000
7.750 1.3527 0.01885 0.01062 -0.0944 0.3306 1.0000
8.000 1.3633 0.01951 0.01123 -0.0918 0.3127 1.0000
8.250 1.3714 0.02035 0.01197 -0.0890 0.2908 1.0000
8.500 1.3767 0.02139 0.01289 -0.0860 0.2643 1.0000
8.750 1.3824 0.02252 0.01394 -0.0832 0.2384 1.0000
9.000 1.3852 0.02392 0.01519 -0.0803 0.2098 1.0000
9.250 1.3853 0.02560 0.01671 -0.0774 0.1777 1.0000
10.000 1.3906 0.03093 0.02182 -0.0705 0.1294 1.0000
10.250 1.3945 0.03272 0.02363 -0.0687 0.1224 1.0000
10.500 1.3995 0.03445 0.02540 -0.0672 0.1142 1.0000
10.750 1.4020 0.03645 0.02742 -0.0657 0.1073 1.0000
11.000 1.4104 0.03800 0.02905 -0.0646 0.0994 1.0000
11.250 1.4141 0.04000 0.03103 -0.0633 0.0737 1.0000
11.500 1.3896 0.04471 0.03540 -0.0608 0.0292 1.0000
11.750 1.3775 0.04851 0.03918 -0.0592 0.0205 1.0000
12.000 1.3763 0.05133 0.04209 -0.0583 0.0178 1.0000
12.250 1.3779 0.05393 0.04483 -0.0575 0.0167 1.0000
12.500 1.3786 0.05670 0.04776 -0.0569 0.0159 1.0000
12.750 1.3779 0.05971 0.05092 -0.0564 0.0150 1.0000
13.000 1.3757 0.06296 0.05432 -0.0560 0.0143 1.0000
13.250 1.3710 0.06659 0.05812 -0.0556 0.0135 1.0000
13.500 1.3710 0.06975 0.06142 -0.0555 0.0129 1.0000
13.750 1.3701 0.07306 0.06489 -0.0555 0.0125 1.0000
14.000 1.3678 0.07662 0.06861 -0.0556 0.0121 1.0000
14.250 1.3643 0.08040 0.07254 -0.0558 0.0118 1.0000
14.500 1.3596 0.08441 0.07671 -0.0561 0.0115 1.0000
14.750 1.3538 0.08868 0.08115 -0.0567 0.0112 1.0000
15.000 1.3469 0.09319 0.08581 -0.0574 0.0110 1.0000
15.250 1.3390 0.09791 0.09069 -0.0583 0.0108 1.0000
15.500 1.3301 0.10285 0.09578 -0.0594 0.0106 1.0000
15.750 1.3202 0.10807 0.10113 -0.0607 0.0104 1.0000
16.000 1.3098 0.11346 0.10666 -0.0622 0.0102 1.0000
16.250 1.2991 0.11896 0.11229 -0.0639 0.0101 1.0000
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Polar data table (+)
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