GOE 604 AIRFOIL (goe604-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 604 AIRFOIL (goe604-il) Reynolds number: 100,000 Max Cl/Cd: 29.83 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe604-il-100000.txt Download as CSV file: xf-goe604-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 604 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.1653 0.10284 0.09870 -0.0574 0.9691 0.1644
-9.500 -0.1489 0.09678 0.09265 -0.0647 0.9571 0.1676
-9.250 -0.2530 0.06719 0.06284 -0.0961 0.9375 0.1098
-9.000 -0.2324 0.06186 0.05742 -0.1006 0.9221 0.1083
-8.750 -0.2351 0.05502 0.05037 -0.1052 0.9037 0.1058
-8.500 -0.3177 0.04283 0.03690 -0.1037 0.8770 0.0990
-8.250 -0.2964 0.03972 0.03351 -0.1043 0.8649 0.1000
-8.000 -0.2745 0.03832 0.03202 -0.1037 0.8491 0.1025
-7.750 -0.2627 0.03617 0.02952 -0.1021 0.8338 0.1045
-7.500 -0.2503 0.03391 0.02681 -0.1003 0.8200 0.1065
-7.250 -0.2328 0.03165 0.02397 -0.0988 0.8092 0.1086
-7.000 -0.2140 0.03026 0.02240 -0.0974 0.7959 0.1114
-6.750 -0.1911 0.02951 0.02158 -0.0965 0.7838 0.1154
-6.500 -0.1669 0.02824 0.01993 -0.0955 0.7742 0.1200
-6.250 -0.1468 0.02722 0.01877 -0.0941 0.7621 0.1240
-6.000 -0.1211 0.02661 0.01805 -0.0934 0.7527 0.1306
-5.750 -0.0992 0.02582 0.01710 -0.0922 0.7418 0.1378
-5.500 -0.0747 0.02527 0.01644 -0.0913 0.7325 0.1470
-5.250 -0.0515 0.02466 0.01585 -0.0902 0.7227 0.1583
-5.000 -0.0281 0.02411 0.01525 -0.0892 0.7139 0.1740
-4.750 -0.0062 0.02365 0.01477 -0.0879 0.7045 0.1953
-4.500 0.0168 0.02328 0.01439 -0.0868 0.6964 0.2204
-4.250 0.0380 0.02317 0.01433 -0.0854 0.6870 0.2427
-4.000 0.0638 0.02306 0.01412 -0.0847 0.6800 0.2644
-3.750 0.0839 0.02306 0.01416 -0.0833 0.6704 0.2815
-3.500 0.1109 0.02290 0.01389 -0.0827 0.6635 0.2996
-3.250 0.1315 0.02292 0.01393 -0.0814 0.6547 0.3157
-3.000 0.1568 0.02279 0.01376 -0.0806 0.6472 0.3343
-2.750 0.1803 0.02276 0.01374 -0.0797 0.6398 0.3544
-2.500 0.2027 0.02269 0.01372 -0.0786 0.6316 0.3745
-2.250 0.2313 0.02241 0.01340 -0.0783 0.6257 0.3964
-2.000 0.2507 0.02239 0.01352 -0.0769 0.6170 0.4199
-1.750 0.2761 0.02209 0.01333 -0.0761 0.6100 0.4526
-1.500 0.2999 0.02187 0.01329 -0.0751 0.6035 0.4962
-1.250 0.3189 0.02171 0.01347 -0.0733 0.5953 0.5560
-1.000 0.3461 0.02123 0.01345 -0.0722 0.5896 0.6681
-0.750 0.3945 0.02132 0.01408 -0.0745 0.5811 0.8471
-0.500 0.4738 0.02169 0.01423 -0.0827 0.5724 0.9355
-0.250 0.5526 0.02215 0.01445 -0.0920 0.5630 0.9750
0.000 0.6306 0.02219 0.01420 -0.1017 0.5542 1.0000
0.250 0.6482 0.02242 0.01424 -0.1003 0.5495 1.0000
0.500 0.6558 0.02298 0.01484 -0.0975 0.5423 1.0000
0.750 0.6730 0.02328 0.01501 -0.0959 0.5369 1.0000
1.000 0.6959 0.02349 0.01502 -0.0950 0.5328 1.0000
1.250 0.7028 0.02426 0.01586 -0.0920 0.5261 1.0000
1.500 0.7187 0.02470 0.01623 -0.0901 0.5205 1.0000
1.750 0.7422 0.02491 0.01627 -0.0893 0.5162 1.0000
2.000 0.7567 0.02556 0.01688 -0.0872 0.5112 1.0000
2.250 0.7653 0.02640 0.01777 -0.0844 0.5053 1.0000
2.500 0.7844 0.02686 0.01816 -0.0830 0.5010 1.0000
2.750 0.8100 0.02715 0.01830 -0.0825 0.4977 1.0000
3.000 0.8199 0.02816 0.01934 -0.0800 0.4933 1.0000
3.250 0.8218 0.02946 0.02074 -0.0765 0.4880 1.0000
3.500 0.8399 0.03006 0.02129 -0.0750 0.4840 1.0000
3.750 0.8657 0.03034 0.02147 -0.0745 0.4808 1.0000
4.000 0.8925 0.03075 0.02176 -0.0743 0.4778 1.0000
4.250 0.8636 0.03350 0.02480 -0.0671 0.4715 1.0000
4.500 0.8668 0.03494 0.02628 -0.0641 0.4674 1.0000
4.750 0.8896 0.03548 0.02676 -0.0634 0.4646 1.0000
5.000 0.9208 0.03570 0.02689 -0.0637 0.4624 1.0000
5.250 0.4271 0.07967 0.07185 -0.0439 0.4695 1.0000
5.500 0.4562 0.08034 0.07245 -0.0435 0.4643 1.0000
5.750 0.4750 0.08224 0.07431 -0.0434 0.4628 1.0000
6.000 0.4802 0.08516 0.07723 -0.0434 0.4625 1.0000
6.250 0.4874 0.08804 0.08009 -0.0434 0.4621 1.0000
6.500 0.4968 0.09090 0.08294 -0.0435 0.4617 1.0000
6.750 0.4960 0.09171 0.08373 -0.0420 0.4462 1.0000
7.000 0.5098 0.09425 0.08625 -0.0421 0.4452 1.0000
7.250 0.5170 0.09733 0.08934 -0.0423 0.4448 1.0000
7.500 0.5131 0.09868 0.09067 -0.0411 0.4301 1.0000
7.750 0.5726 0.09326 0.08510 -0.0380 0.3979 1.0000
8.000 0.6582 0.08765 0.07936 -0.0361 0.3930 1.0000
8.250 0.5927 0.09684 0.08866 -0.0372 0.3833 1.0000
8.500 0.6337 0.09567 0.08746 -0.0362 0.3784 1.0000
8.750 0.6174 0.10004 0.09186 -0.0364 0.3686 1.0000
9.000 0.6473 0.09979 0.09158 -0.0356 0.3628 1.0000
9.250 0.6995 0.09720 0.08894 -0.0341 0.3602 1.0000
9.500 0.6715 0.10267 0.09448 -0.0347 0.3468 1.0000
9.750 0.6716 0.10557 0.09739 -0.0346 0.3375 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 604 AIRFOIL (goe604-il)