GOE 625 AIRFOIL (goe625-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 625 AIRFOIL (goe625-il) Reynolds number: 100,000 Max Cl/Cd: 33.81 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe625-il-100000.txt Download as CSV file: xf-goe625-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 625 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.1389 0.13423 0.12933 -0.0615 0.9692 0.1437
-10.750 -0.1346 0.13113 0.12621 -0.0677 0.9657 0.1498
-10.500 -0.1588 0.12849 0.12363 -0.0723 0.9571 0.1518
-10.250 -0.1059 0.12287 0.11795 -0.0732 0.9544 0.1543
-10.000 -0.0755 0.11896 0.11400 -0.0770 0.9514 0.1592
-9.750 -0.1036 0.11769 0.11277 -0.0816 0.9415 0.1657
-9.500 -0.0870 0.11228 0.10735 -0.0855 0.9373 0.1678
-9.250 -0.0380 0.10773 0.10273 -0.0879 0.9354 0.1714
-9.000 0.0327 0.09831 0.09353 -0.1011 0.9026 0.1841
-8.750 0.0934 0.09292 0.08807 -0.1032 0.9007 0.1881
-8.500 0.1144 0.08895 0.08406 -0.1066 0.8977 0.1963
-8.250 0.0224 0.08896 0.08416 -0.1077 0.8808 0.2010
-8.000 0.1131 0.08230 0.07741 -0.1081 0.8808 0.2051
-7.750 0.1291 0.07999 0.07510 -0.1072 0.8715 0.2103
-7.500 0.0487 0.07799 0.07316 -0.1102 0.8594 0.2198
-7.250 0.0044 0.08311 0.07810 -0.1059 0.8722 0.2184
-7.000 0.0047 0.07887 0.07386 -0.1072 0.8635 0.2227
-6.750 0.0621 0.07591 0.07083 -0.1080 0.8614 0.2298
-6.500 -0.0121 0.07316 0.06803 -0.1101 0.8426 0.2411
-6.250 0.0460 0.07006 0.06495 -0.1092 0.8408 0.2451
-6.000 -0.0658 0.05102 0.04456 -0.1177 0.8216 0.1678
-5.750 -0.0652 0.04864 0.04182 -0.1149 0.8075 0.1656
-5.500 -0.0340 0.04498 0.03766 -0.1170 0.8032 0.1663
-5.250 -0.0225 0.04293 0.03538 -0.1150 0.7906 0.1668
-5.000 0.0151 0.04025 0.03268 -0.1169 0.7854 0.1691
-4.750 0.0410 0.03882 0.03118 -0.1166 0.7757 0.1716
-4.500 0.0720 0.03686 0.02897 -0.1171 0.7673 0.1738
-4.250 0.1132 0.03452 0.02626 -0.1191 0.7630 0.1763
-4.000 0.1306 0.03350 0.02501 -0.1173 0.7485 0.1789
-3.750 0.1723 0.03172 0.02274 -0.1191 0.7435 0.1825
-3.500 0.1914 0.03063 0.02154 -0.1176 0.7290 0.1849
-3.250 0.2209 0.02957 0.02041 -0.1176 0.7186 0.1885
-3.000 0.2532 0.02862 0.01931 -0.1180 0.7087 0.1939
-2.750 0.2794 0.02802 0.01847 -0.1174 0.6968 0.1993
-2.500 0.3143 0.02680 0.01712 -0.1183 0.6884 0.2053
-2.250 0.3376 0.02640 0.01668 -0.1173 0.6761 0.2119
-2.000 0.3734 0.02550 0.01556 -0.1183 0.6685 0.2226
-1.750 0.3942 0.02528 0.01542 -0.1171 0.6570 0.2329
-1.500 0.4280 0.02455 0.01460 -0.1179 0.6501 0.2477
-1.250 0.4495 0.02448 0.01453 -0.1168 0.6405 0.2623
-1.000 0.4791 0.02414 0.01416 -0.1170 0.6334 0.2808
-0.750 0.5070 0.02390 0.01399 -0.1170 0.6269 0.2998
-0.500 0.5284 0.02391 0.01413 -0.1160 0.6188 0.3182
-0.250 0.5600 0.02360 0.01384 -0.1165 0.6132 0.3447
0.000 0.5826 0.02360 0.01406 -0.1157 0.6068 0.3775
0.250 0.6035 0.02349 0.01435 -0.1147 0.6004 0.4438
0.500 0.6435 0.02231 0.01472 -0.1144 0.5957 0.9258
0.750 0.7305 0.02268 0.01472 -0.1258 0.5899 1.0000
1.000 0.7403 0.02330 0.01528 -0.1230 0.5841 1.0000
1.250 0.7601 0.02365 0.01548 -0.1218 0.5791 1.0000
1.500 0.7883 0.02385 0.01545 -0.1219 0.5752 1.0000
1.750 0.8061 0.02445 0.01595 -0.1205 0.5711 1.0000
2.000 0.8142 0.02531 0.01683 -0.1176 0.5662 1.0000
2.250 0.8326 0.02588 0.01732 -0.1162 0.5617 1.0000
2.500 0.8598 0.02619 0.01748 -0.1162 0.5578 1.0000
2.750 0.8944 0.02645 0.01752 -0.1174 0.5546 1.0000
3.000 0.8931 0.02776 0.01897 -0.1132 0.5497 1.0000
3.250 0.9044 0.02870 0.01992 -0.1109 0.5449 1.0000
3.500 0.9295 0.02912 0.02025 -0.1106 0.5407 1.0000
3.750 0.9670 0.02916 0.02009 -0.1121 0.5370 1.0000
4.000 0.9710 0.03041 0.02141 -0.1089 0.5318 1.0000
4.250 0.9768 0.03151 0.02257 -0.1059 0.5259 1.0000
4.500 1.0071 0.03168 0.02262 -0.1063 0.5213 1.0000
4.750 1.0507 0.03152 0.02226 -0.1087 0.5177 1.0000
5.000 1.0388 0.03349 0.02440 -0.1033 0.5124 1.0000
5.250 1.0318 0.03528 0.02630 -0.0989 0.5068 1.0000
5.500 1.0621 0.03551 0.02645 -0.0993 0.5027 1.0000
5.750 1.1109 0.03510 0.02587 -0.1023 0.4994 1.0000
6.000 1.0815 0.03800 0.02897 -0.0951 0.4937 1.0000
6.250 0.9812 0.04500 0.03628 -0.0810 0.4841 1.0000
6.500 1.0738 0.04139 0.03245 -0.0875 0.4830 1.0000
6.750 1.1533 0.03959 0.03045 -0.0941 0.4805 1.0000
7.000 0.8964 0.06048 0.05198 -0.0710 0.4588 1.0000
7.250 0.9603 0.05669 0.04808 -0.0721 0.4589 1.0000
7.500 1.0351 0.05251 0.04377 -0.0745 0.4589 1.0000
7.750 1.1402 0.04715 0.03821 -0.0805 0.4591 1.0000
10.000 0.7071 0.12423 0.11622 -0.0669 0.3607 1.0000
10.250 0.6972 0.12975 0.12179 -0.0680 0.3610 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 625 AIRFOIL (goe625-il)