GOE 500 AIRFOIL (goe500-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 500 AIRFOIL (goe500-il) Reynolds number: 500,000 Max Cl/Cd: 120.62 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe500-il-500000-n5.txt Download as CSV file: xf-goe500-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 500 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.4746 0.02638 0.02301 -0.1558 0.9300 0.0157
-10.750 -0.4546 0.02327 0.01957 -0.1597 0.9244 0.0164
-10.500 -0.4312 0.02145 0.01753 -0.1618 0.9203 0.0170
-10.250 -0.4052 0.02046 0.01643 -0.1629 0.9172 0.0177
-10.000 -0.3801 0.01971 0.01556 -0.1635 0.9136 0.0185
-9.750 -0.3548 0.01893 0.01464 -0.1640 0.9097 0.0195
-9.500 -0.3290 0.01812 0.01364 -0.1645 0.9062 0.0205
-9.250 -0.3027 0.01732 0.01269 -0.1651 0.9031 0.0214
-9.000 -0.2760 0.01680 0.01209 -0.1655 0.9000 0.0223
-8.750 -0.2497 0.01634 0.01157 -0.1656 0.8962 0.0232
-8.500 -0.2230 0.01586 0.01096 -0.1658 0.8925 0.0242
-8.250 -0.1958 0.01538 0.01035 -0.1660 0.8891 0.0253
-8.000 -0.1682 0.01494 0.00975 -0.1663 0.8862 0.0261
-7.750 -0.1420 0.01424 0.00894 -0.1664 0.8822 0.0270
-7.500 -0.1152 0.01376 0.00841 -0.1666 0.8781 0.0281
-7.250 -0.0877 0.01339 0.00795 -0.1667 0.8744 0.0291
-7.000 -0.0599 0.01301 0.00748 -0.1668 0.8712 0.0301
-6.750 -0.0327 0.01265 0.00703 -0.1669 0.8668 0.0312
-6.500 -0.0053 0.01230 0.00659 -0.1669 0.8624 0.0321
-6.250 0.0225 0.01196 0.00613 -0.1669 0.8584 0.0326
-6.000 0.0501 0.01145 0.00554 -0.1670 0.8544 0.0338
-5.750 0.0775 0.01106 0.00510 -0.1671 0.8491 0.0350
-5.500 0.1053 0.01074 0.00472 -0.1671 0.8439 0.0360
-5.250 0.1331 0.01046 0.00437 -0.1670 0.8383 0.0370
-5.000 0.1606 0.01021 0.00405 -0.1669 0.8310 0.0382
-4.750 0.1885 0.00999 0.00374 -0.1669 0.8240 0.0395
-4.500 0.2161 0.00979 0.00348 -0.1667 0.8156 0.0407
-4.250 0.2437 0.00953 0.00315 -0.1667 0.8067 0.0428
-4.000 0.2714 0.00933 0.00289 -0.1665 0.7971 0.0454
-3.750 0.2990 0.00917 0.00268 -0.1664 0.7875 0.0486
-3.500 0.3268 0.00902 0.00247 -0.1663 0.7795 0.0537
-3.250 0.3545 0.00887 0.00232 -0.1662 0.7711 0.0626
-3.000 0.3821 0.00877 0.00224 -0.1661 0.7630 0.0773
-2.750 0.4097 0.00872 0.00217 -0.1660 0.7539 0.0882
-2.500 0.4373 0.00870 0.00210 -0.1658 0.7452 0.0960
-2.250 0.4647 0.00869 0.00202 -0.1656 0.7368 0.1005
-2.000 0.4924 0.00863 0.00196 -0.1655 0.7289 0.1058
-1.500 0.5473 0.00863 0.00185 -0.1652 0.7128 0.1138
-1.250 0.5745 0.00859 0.00181 -0.1650 0.7047 0.1230
-1.000 0.6019 0.00857 0.00178 -0.1649 0.6961 0.1318
-0.750 0.6290 0.00856 0.00176 -0.1647 0.6883 0.1394
-0.500 0.6564 0.00854 0.00175 -0.1645 0.6803 0.1496
0.000 0.7107 0.00845 0.00180 -0.1643 0.6646 0.2158
0.250 0.7376 0.00846 0.00185 -0.1640 0.6566 0.2472
0.500 0.7646 0.00848 0.00189 -0.1638 0.6476 0.2713
0.750 0.7913 0.00851 0.00195 -0.1636 0.6388 0.2922
1.000 0.8179 0.00856 0.00202 -0.1633 0.6295 0.3119
1.250 0.8447 0.00860 0.00209 -0.1630 0.6207 0.3324
1.500 0.8709 0.00866 0.00218 -0.1627 0.6116 0.3590
1.750 0.8974 0.00869 0.00228 -0.1624 0.6019 0.3878
2.000 0.9235 0.00874 0.00239 -0.1620 0.5922 0.4252
2.250 0.9490 0.00880 0.00253 -0.1616 0.5813 0.4713
2.500 0.9748 0.00885 0.00267 -0.1612 0.5699 0.5163
2.750 1.0001 0.00893 0.00282 -0.1607 0.5579 0.5592
3.250 1.0486 0.00909 0.00315 -0.1593 0.5285 0.6698
3.750 1.0952 0.00908 0.00349 -0.1574 0.4909 1.0000
4.000 1.1194 0.00932 0.00367 -0.1567 0.4771 1.0000
4.250 1.1426 0.00961 0.00388 -0.1558 0.4606 1.0000
4.500 1.1654 0.00992 0.00410 -0.1549 0.4439 1.0000
4.750 1.1893 0.01017 0.00432 -0.1542 0.4317 1.0000
5.000 1.2124 0.01046 0.00456 -0.1533 0.4178 1.0000
5.250 1.2340 0.01081 0.00483 -0.1522 0.3990 1.0000
5.500 1.2564 0.01112 0.00510 -0.1512 0.3833 1.0000
5.750 1.2782 0.01145 0.00537 -0.1502 0.3663 1.0000
6.000 1.2987 0.01184 0.00568 -0.1489 0.3413 1.0000
6.250 1.3155 0.01243 0.00608 -0.1470 0.3026 1.0000
6.500 1.3257 0.01341 0.00673 -0.1441 0.2455 1.0000
6.750 1.3343 0.01437 0.00742 -0.1409 0.2000 1.0000
7.000 1.3430 0.01532 0.00813 -0.1377 0.1553 1.0000
7.250 1.3496 0.01640 0.00894 -0.1344 0.1113 1.0000
7.500 1.3570 0.01745 0.00978 -0.1312 0.0751 1.0000
7.750 1.3665 0.01839 0.01059 -0.1284 0.0533 1.0000
8.000 1.3803 0.01909 0.01128 -0.1263 0.0458 1.0000
8.250 1.3952 0.01973 0.01194 -0.1245 0.0417 1.0000
8.500 1.4096 0.02041 0.01263 -0.1226 0.0385 1.0000
8.750 1.4229 0.02118 0.01342 -0.1206 0.0356 1.0000
9.000 1.4375 0.02187 0.01416 -0.1188 0.0338 1.0000
9.250 1.4518 0.02259 0.01494 -0.1171 0.0319 1.0000
9.500 1.4648 0.02342 0.01581 -0.1152 0.0300 1.0000
9.750 1.4760 0.02439 0.01681 -0.1132 0.0280 1.0000
10.000 1.4889 0.02526 0.01773 -0.1115 0.0265 1.0000
10.250 1.5020 0.02614 0.01867 -0.1099 0.0250 1.0000
10.500 1.5136 0.02713 0.01971 -0.1081 0.0232 1.0000
10.750 1.5230 0.02833 0.02094 -0.1063 0.0214 1.0000
11.000 1.5348 0.02938 0.02206 -0.1047 0.0201 1.0000
11.250 1.5453 0.03054 0.02327 -0.1031 0.0185 1.0000
11.500 1.5542 0.03187 0.02463 -0.1014 0.0171 1.0000
11.750 1.5622 0.03332 0.02614 -0.0997 0.0161 1.0000
12.000 1.5705 0.03478 0.02768 -0.0982 0.0151 1.0000
12.250 1.5779 0.03635 0.02932 -0.0967 0.0142 1.0000
12.500 1.5842 0.03809 0.03111 -0.0953 0.0133 1.0000
12.750 1.5887 0.04006 0.03314 -0.0938 0.0125 1.0000
13.000 1.5946 0.04195 0.03513 -0.0926 0.0120 1.0000
13.250 1.5995 0.04399 0.03727 -0.0914 0.0114 1.0000
13.500 1.6040 0.04616 0.03952 -0.0904 0.0109 1.0000
13.750 1.6076 0.04850 0.04194 -0.0894 0.0105 1.0000
14.000 1.6100 0.05105 0.04457 -0.0886 0.0101 1.0000
14.250 1.6104 0.05391 0.04751 -0.0878 0.0097 1.0000
14.500 1.6109 0.05685 0.05057 -0.0873 0.0095 1.0000
14.750 1.6128 0.05969 0.05353 -0.0868 0.0092 1.0000
15.000 1.6140 0.06270 0.05664 -0.0865 0.0089 1.0000
15.250 1.6143 0.06589 0.05994 -0.0864 0.0086 1.0000
15.500 1.6147 0.06914 0.06330 -0.0863 0.0083 1.0000
15.750 1.6145 0.07257 0.06683 -0.0865 0.0081 1.0000
16.000 1.6140 0.07610 0.07047 -0.0868 0.0078 1.0000
16.250 1.6117 0.07998 0.07445 -0.0872 0.0077 1.0000
16.500 1.6086 0.08407 0.07864 -0.0879 0.0075 1.0000
16.750 1.6036 0.08855 0.08323 -0.0887 0.0073 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 500 AIRFOIL (goe500-il)