GOE 55 AIRFOIL (goe55-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 55 AIRFOIL (goe55-il) Reynolds number: 50,000 Max Cl/Cd: 26.29 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe55-il-50000.txt Download as CSV file: xf-goe55-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 55 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.6256 0.12123 0.11417 0.0423 1.0000 0.1673
-8.500 -0.6227 0.11834 0.11132 0.0410 1.0000 0.1749
-8.250 -0.6408 0.11822 0.11133 0.0349 1.0000 0.1778
-8.000 -0.6105 0.11080 0.10386 0.0395 1.0000 0.1859
-7.750 -0.6199 0.10933 0.10249 0.0351 1.0000 0.1926
-7.500 -0.6048 0.10409 0.09726 0.0366 1.0000 0.1983
-7.250 -0.6049 0.10170 0.09493 0.0321 1.0000 0.2079
-6.750 -0.5890 0.09400 0.08729 0.0275 1.0000 0.2258
-6.500 -0.5807 0.09108 0.08438 0.0228 1.0000 0.2404
-6.000 -0.5570 0.08343 0.07679 0.0223 1.0000 0.2722
-5.750 -0.5441 0.07947 0.07287 0.0236 1.0000 0.2889
-5.500 -0.5308 0.07572 0.06916 0.0257 1.0000 0.3072
-5.250 -0.5191 0.07280 0.06626 0.0237 1.0000 0.3365
-5.000 -0.5057 0.06924 0.06278 0.0285 1.0000 0.3600
-4.750 -0.4963 0.06639 0.06000 0.0302 1.0000 0.4015
-4.500 -0.4876 0.06366 0.05738 0.0359 1.0000 0.4492
-4.250 -0.4779 0.06093 0.05473 0.0443 1.0000 0.5035
-3.750 -0.1866 0.04154 0.03490 0.0362 1.0000 1.0000
-3.500 -0.1878 0.04007 0.03351 0.0382 1.0000 0.9936
-3.250 -0.2457 0.04133 0.03499 0.0520 1.0000 0.9445
-3.000 -0.2184 0.03551 0.02676 -0.0247 1.0000 0.2171
-2.750 -0.1767 0.03232 0.02286 -0.0276 1.0000 0.1895
-2.500 -0.1397 0.02991 0.01981 -0.0290 1.0000 0.1770
-2.250 -0.1083 0.02780 0.01750 -0.0295 1.0000 0.1749
-2.000 -0.0763 0.02617 0.01552 -0.0298 1.0000 0.1769
-1.750 -0.0453 0.02457 0.01373 -0.0299 1.0000 0.1788
-1.500 -0.0158 0.02305 0.01219 -0.0298 1.0000 0.1818
-1.250 0.0135 0.02185 0.01096 -0.0295 1.0000 0.1891
-1.000 0.0420 0.02072 0.00987 -0.0289 1.0000 0.2017
-0.750 0.0715 0.01964 0.00885 -0.0285 1.0000 0.2163
-0.500 0.1003 0.01855 0.00800 -0.0281 1.0000 0.2455
-0.250 0.1295 0.01713 0.00721 -0.0281 1.0000 0.3109
0.000 0.1516 0.01418 0.00624 -0.0252 1.0000 1.0000
0.250 0.1845 0.01433 0.00641 -0.0272 1.0000 1.0000
0.500 0.2525 0.01532 0.00668 -0.0344 0.7913 1.0000
0.750 0.2759 0.01616 0.00694 -0.0323 0.7338 1.0000
1.000 0.3023 0.01690 0.00730 -0.0314 0.6973 1.0000
1.250 0.3296 0.01759 0.00772 -0.0308 0.6704 1.0000
1.500 0.3582 0.01824 0.00823 -0.0307 0.6467 1.0000
1.750 0.3863 0.01890 0.00874 -0.0304 0.6258 1.0000
2.000 0.4136 0.01957 0.00927 -0.0299 0.6055 1.0000
2.250 0.4412 0.02024 0.00989 -0.0296 0.5835 1.0000
2.500 0.4679 0.02094 0.01050 -0.0289 0.5624 1.0000
2.750 0.4943 0.02171 0.01116 -0.0282 0.5427 1.0000
3.000 0.5219 0.02254 0.01202 -0.0280 0.5216 1.0000
3.250 0.5489 0.02344 0.01292 -0.0276 0.5006 1.0000
3.500 0.5754 0.02442 0.01386 -0.0269 0.4803 1.0000
3.750 0.6024 0.02557 0.01507 -0.0267 0.4589 1.0000
4.000 0.6293 0.02681 0.01642 -0.0264 0.4368 1.0000
4.250 0.6551 0.02806 0.01766 -0.0257 0.4162 1.0000
4.500 0.6821 0.02967 0.01944 -0.0257 0.3956 1.0000
4.750 0.7078 0.03101 0.02094 -0.0253 0.3723 1.0000
5.000 0.7317 0.03181 0.02175 -0.0241 0.3452 1.0000
5.250 0.7543 0.03160 0.02139 -0.0222 0.3118 1.0000
5.500 0.7779 0.03135 0.02105 -0.0206 0.2792 1.0000
5.750 0.8024 0.03121 0.02083 -0.0190 0.2447 1.0000
6.000 0.8275 0.03147 0.02100 -0.0176 0.2048 1.0000
6.250 0.8526 0.03260 0.02200 -0.0167 0.1800 1.0000
6.500 0.8776 0.03498 0.02454 -0.0164 0.1662 1.0000
6.750 0.9023 0.03727 0.02684 -0.0159 0.1575 1.0000
7.000 0.9236 0.04189 0.03223 -0.0167 0.1514 1.0000
7.250 0.9450 0.04495 0.03553 -0.0167 0.1453 1.0000
7.500 0.9637 0.04885 0.03968 -0.0170 0.1412 1.0000
7.750 0.9737 0.05570 0.04723 -0.0191 0.1406 1.0000
8.000 0.9785 0.06311 0.05512 -0.0217 0.1413 1.0000
8.250 0.9838 0.06981 0.06212 -0.0238 0.1423 1.0000
8.500 0.9006 0.09812 0.09086 -0.0523 0.1671 1.0000
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Polar data table (+)
Polar graphs
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