GOE 81 AIRFOIL (goe81-il) Xfoil prediction polar at RE=100,000 Ncrit=5
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Airfoil: GOE 81 AIRFOIL (goe81-il) Reynolds number: 100,000 Max Cl/Cd: 65.07 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe81-il-100000-n5.txt Download as CSV file: xf-goe81-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 81 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.2608 0.11274 0.10801 -0.0339 1.0000 0.0334
-8.500 -0.2729 0.11238 0.10777 -0.0319 1.0000 0.0335
-8.250 -0.2814 0.11172 0.10721 -0.0315 0.9972 0.0335
-8.000 -0.2612 0.10459 0.10008 -0.0298 0.9974 0.0342
-7.750 -0.2427 0.10034 0.09582 -0.0323 0.9920 0.0354
-7.500 -0.2244 0.09681 0.09228 -0.0364 0.9838 0.0370
-7.250 -0.2041 0.09326 0.08872 -0.0417 0.9745 0.0387
-7.000 -0.1800 0.08994 0.08538 -0.0493 0.9639 0.0404
-6.750 -0.1450 0.08737 0.08272 -0.0635 0.9517 0.0412
-6.500 -0.1226 0.08297 0.07831 -0.0682 0.9436 0.0417
-6.250 -0.1098 0.07839 0.07376 -0.0658 0.9388 0.0429
-6.000 -0.0870 0.07493 0.07025 -0.0691 0.9316 0.0451
-5.750 -0.0594 0.07154 0.06681 -0.0748 0.9233 0.0475
-5.500 -0.0135 0.06886 0.06394 -0.0876 0.9132 0.0506
-5.250 0.0228 0.06503 0.05997 -0.0952 0.9063 0.0513
-5.000 0.0335 0.06129 0.05629 -0.0934 0.8985 0.0525
-4.750 0.0569 0.05835 0.05330 -0.0953 0.8914 0.0554
-4.500 0.1024 0.05635 0.05104 -0.1040 0.8817 0.0613
-4.250 0.1337 0.03522 0.03019 -0.0995 0.8448 0.0635
-4.000 0.1505 0.03261 0.02755 -0.0997 0.8346 0.0655
-3.750 0.1759 0.03000 0.02482 -0.1019 0.8263 0.0689
-3.500 0.2168 0.02802 0.02247 -0.1080 0.8160 0.0751
-3.250 0.2309 0.02518 0.01969 -0.1072 0.8070 0.0776
-3.000 0.2558 0.02309 0.01749 -0.1084 0.7985 0.0823
-2.750 0.2928 0.02152 0.01555 -0.1121 0.7893 0.0894
-2.500 0.3141 0.01919 0.01321 -0.1125 0.7828 0.0921
-2.250 0.3446 0.01821 0.01199 -0.1140 0.7743 0.1016
-2.000 0.3716 0.01615 0.00980 -0.1153 0.7677 0.1059
-1.750 0.3970 0.01488 0.00844 -0.1159 0.7597 0.1124
-1.500 0.4247 0.01361 0.00702 -0.1168 0.7525 0.1261
-1.000 0.5233 0.02611 0.01797 -0.1269 0.7559 0.0626
-0.750 0.5514 0.02466 0.01642 -0.1272 0.7483 0.0590
-0.500 0.5807 0.02344 0.01498 -0.1273 0.7394 0.0561
-0.250 0.6114 0.02229 0.01346 -0.1273 0.7316 0.0532
0.000 0.6407 0.02147 0.01230 -0.1271 0.7227 0.0520
0.250 0.6683 0.02077 0.01152 -0.1270 0.7140 0.0536
0.500 0.6964 0.02016 0.01076 -0.1268 0.7053 0.0548
0.750 0.7240 0.01959 0.01003 -0.1264 0.6956 0.0543
1.000 0.7521 0.01906 0.00931 -0.1260 0.6871 0.0539
1.250 0.7788 0.01866 0.00885 -0.1255 0.6766 0.0539
1.500 0.8056 0.01831 0.00843 -0.1251 0.6666 0.0543
1.750 0.8323 0.01800 0.00805 -0.1246 0.6568 0.0549
2.000 0.8578 0.01780 0.00783 -0.1239 0.6452 0.0562
2.250 0.8840 0.01769 0.00764 -0.1233 0.6333 0.0597
2.500 0.9104 0.01760 0.00745 -0.1228 0.6211 0.0625
2.750 0.9367 0.01754 0.00730 -0.1223 0.6081 0.0650
3.000 0.9624 0.01756 0.00727 -0.1217 0.5935 0.0688
3.500 1.0136 0.01764 0.00735 -0.1206 0.5635 0.1005
3.750 1.0351 0.01642 0.00752 -0.1193 0.5492 1.0000
4.000 1.0599 0.01665 0.00762 -0.1187 0.5341 1.0000
4.250 1.0845 0.01690 0.00776 -0.1180 0.5189 1.0000
4.500 1.1089 0.01717 0.00795 -0.1173 0.5038 1.0000
4.750 1.1331 0.01747 0.00815 -0.1166 0.4894 1.0000
5.000 1.1571 0.01779 0.00839 -0.1158 0.4761 1.0000
5.250 1.1810 0.01815 0.00869 -0.1151 0.4640 1.0000
5.500 1.2047 0.01854 0.00899 -0.1144 0.4527 1.0000
5.750 1.2281 0.01896 0.00938 -0.1137 0.4407 1.0000
6.000 1.2508 0.01940 0.00981 -0.1128 0.4277 1.0000
6.250 1.2730 0.01987 0.01024 -0.1119 0.4143 1.0000
6.500 1.2948 0.02036 0.01070 -0.1110 0.4012 1.0000
6.750 1.3161 0.02086 0.01119 -0.1100 0.3877 1.0000
7.000 1.3373 0.02136 0.01173 -0.1090 0.3751 1.0000
7.250 1.3587 0.02184 0.01230 -0.1081 0.3633 1.0000
7.500 1.3797 0.02234 0.01287 -0.1071 0.3526 1.0000
7.750 1.4000 0.02287 0.01347 -0.1060 0.3421 1.0000
8.000 1.4199 0.02339 0.01411 -0.1049 0.3301 1.0000
8.250 1.4389 0.02395 0.01478 -0.1036 0.3176 1.0000
8.500 1.4566 0.02455 0.01546 -0.1022 0.3036 1.0000
8.750 1.4727 0.02522 0.01621 -0.1006 0.2883 1.0000
9.000 1.4873 0.02595 0.01699 -0.0987 0.2722 1.0000
9.250 1.4994 0.02680 0.01785 -0.0966 0.2544 1.0000
9.500 1.5079 0.02779 0.01880 -0.0941 0.2354 1.0000
9.750 1.5146 0.02887 0.01986 -0.0913 0.2160 1.0000
10.000 1.5201 0.03010 0.02108 -0.0886 0.1973 1.0000
10.250 1.5250 0.03146 0.02246 -0.0861 0.1810 1.0000
10.500 1.5298 0.03292 0.02393 -0.0838 0.1675 1.0000
10.750 1.5332 0.03456 0.02559 -0.0815 0.1552 1.0000
11.000 1.5351 0.03639 0.02744 -0.0794 0.1443 1.0000
11.250 1.5350 0.03848 0.02953 -0.0774 0.1354 1.0000
11.500 1.5382 0.04040 0.03157 -0.0757 0.1276 1.0000
11.750 1.5368 0.04281 0.03405 -0.0741 0.1210 1.0000
12.000 1.5402 0.04488 0.03629 -0.0728 0.1134 1.0000
12.250 1.5377 0.04760 0.03909 -0.0715 0.1063 1.0000
12.500 1.5406 0.04991 0.04160 -0.0706 0.0945 1.0000
12.750 1.5410 0.05254 0.04438 -0.0698 0.0763 1.0000
13.000 1.5313 0.05642 0.04818 -0.0692 0.0657 1.0000
13.250 1.5190 0.06076 0.05248 -0.0689 0.0591 1.0000
13.500 1.5086 0.06504 0.05685 -0.0687 0.0532 1.0000
13.750 1.4951 0.06988 0.06177 -0.0689 0.0490 1.0000
14.000 1.4874 0.07414 0.06624 -0.0691 0.0434 1.0000
14.500 1.4670 0.08385 0.07626 -0.0705 0.0344 1.0000
14.750 1.4552 0.08919 0.08172 -0.0716 0.0317 1.0000
15.000 1.4441 0.09461 0.08729 -0.0730 0.0296 1.0000
15.250 1.4323 0.10031 0.09312 -0.0747 0.0282 1.0000
15.500 1.4199 0.10623 0.09916 -0.0766 0.0271 1.0000
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Polar data table (+)
Polar graphs
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