GOE 411 AIRFOIL (goe411-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 411 AIRFOIL (goe411-il) Reynolds number: 50,000 Max Cl/Cd: 30.14 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe411-il-50000.txt Download as CSV file: xf-goe411-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 411 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
------ -------- --------- --------- -------- -------- --------
-12.250 -0.6232 0.10929 0.10231 -0.0355 1.0000 0.0973
-12.000 -0.6194 0.10489 0.09793 -0.0356 1.0000 0.0953
-11.750 -0.6354 0.09693 0.09003 -0.0395 1.0000 0.0918
-11.500 -0.6819 0.08617 0.07931 -0.0468 1.0000 0.0890
-11.250 -0.7293 0.07878 0.07191 -0.0507 1.0000 0.0876
-11.000 -0.7717 0.07407 0.06704 -0.0507 1.0000 0.0868
-10.750 -0.8119 0.07098 0.06376 -0.0473 1.0000 0.0861
-10.500 -0.8476 0.06920 0.06174 -0.0417 1.0000 0.0853
-10.250 -0.8500 0.06506 0.05747 -0.0394 1.0000 0.0840
-10.000 -0.8558 0.06148 0.05373 -0.0365 1.0000 0.0828
-9.750 -0.8632 0.05812 0.05014 -0.0330 1.0000 0.0817
-9.500 -0.8686 0.05481 0.04653 -0.0294 1.0000 0.0805
-9.250 -0.8701 0.05165 0.04302 -0.0259 1.0000 0.0792
-9.000 -0.8672 0.04844 0.03941 -0.0225 1.0000 0.0783
-8.750 -0.8577 0.04544 0.03594 -0.0198 1.0000 0.0778
-8.500 -0.8423 0.04278 0.03294 -0.0177 1.0000 0.0792
-8.250 -0.8217 0.04030 0.03014 -0.0163 1.0000 0.0817
-8.000 -0.7923 0.03796 0.02745 -0.0158 1.0000 0.0842
-7.750 -0.7512 0.03538 0.02487 -0.0169 1.0000 0.0876
-7.500 -0.7320 0.03371 0.02321 -0.0150 1.0000 0.0917
-7.250 -0.7203 0.03231 0.02169 -0.0120 1.0000 0.0965
-7.000 -0.7153 0.03084 0.02030 -0.0083 1.0000 0.1029
-6.750 -0.7113 0.02938 0.01900 -0.0044 1.0000 0.1161
-6.500 -0.7087 0.02778 0.01770 -0.0004 1.0000 0.1466
-6.250 -0.7142 0.02572 0.01647 0.0049 1.0000 0.2403
-6.000 -0.7213 0.02453 0.01602 0.0111 1.0000 0.3514
-5.750 -0.7238 0.02415 0.01589 0.0173 1.0000 0.4224
-5.500 -0.7254 0.02408 0.01596 0.0236 1.0000 0.4780
-5.250 -0.7196 0.02401 0.01589 0.0285 1.0000 0.5189
-5.000 -0.7128 0.02414 0.01610 0.0337 1.0000 0.5641
-4.750 -0.6991 0.02402 0.01592 0.0374 1.0000 0.6000
-4.500 -0.6840 0.02431 0.01627 0.0417 1.0000 0.6441
-4.250 -0.6563 0.02583 0.01788 0.0460 1.0000 0.6968
-4.000 -0.6230 0.02671 0.01861 0.0477 1.0000 0.7330
-3.750 -0.5867 0.02723 0.01889 0.0478 1.0000 0.7614
-3.500 -0.5460 0.02784 0.01921 0.0470 1.0000 0.7886
-3.250 -0.5011 0.02850 0.01960 0.0453 1.0000 0.8151
-3.000 -0.4105 0.03004 0.02072 0.0362 1.0000 0.8388
-2.750 -0.3311 0.03088 0.02119 0.0281 1.0000 0.8633
-2.500 -0.2772 0.03102 0.02112 0.0233 1.0000 0.8874
-2.250 -0.2154 0.03093 0.02081 0.0165 1.0000 0.9088
-2.000 -0.1653 0.03059 0.02033 0.0112 1.0000 0.9301
-1.750 -0.1084 0.03005 0.01966 0.0043 1.0000 0.9501
-1.500 -0.0542 0.02942 0.01894 -0.0026 1.0000 0.9707
-1.250 0.0116 0.02845 0.01790 -0.0121 1.0000 0.9907
-1.000 0.0434 0.02773 0.01714 -0.0157 1.0000 1.0000
-0.750 0.0368 0.02737 0.01682 -0.0123 1.0000 1.0000
-0.500 0.0269 0.02707 0.01656 -0.0085 1.0000 1.0000
-0.250 0.0142 0.02688 0.01638 -0.0044 1.0000 1.0000
0.000 0.0000 0.02681 0.01632 0.0000 1.0000 1.0000
0.250 -0.0142 0.02688 0.01638 0.0044 1.0000 1.0000
0.500 -0.0269 0.02707 0.01656 0.0086 1.0000 1.0000
0.750 -0.0368 0.02736 0.01682 0.0124 1.0000 1.0000
1.000 -0.0434 0.02772 0.01714 0.0157 1.0000 1.0000
1.250 -0.0119 0.02845 0.01789 0.0121 0.9907 1.0000
1.500 0.0537 0.02941 0.01892 0.0027 0.9708 1.0000
1.750 0.1083 0.03005 0.01966 -0.0043 0.9502 1.0000
2.000 0.1654 0.03058 0.02032 -0.0112 0.9301 1.0000
2.250 0.2153 0.03093 0.02080 -0.0165 0.9089 1.0000
2.500 0.2765 0.03103 0.02113 -0.0232 0.8875 1.0000
2.750 0.3316 0.03086 0.02118 -0.0282 0.8632 1.0000
3.000 0.4105 0.03003 0.02071 -0.0362 0.8389 1.0000
3.250 0.5010 0.02850 0.01961 -0.0453 0.8151 1.0000
3.500 0.5460 0.02783 0.01920 -0.0470 0.7886 1.0000
3.750 0.5869 0.02722 0.01888 -0.0478 0.7613 1.0000
4.000 0.6233 0.02670 0.01860 -0.0477 0.7329 1.0000
4.250 0.6560 0.02575 0.01780 -0.0458 0.6954 1.0000
4.500 0.6842 0.02431 0.01628 -0.0417 0.6442 1.0000
4.750 0.6987 0.02401 0.01590 -0.0373 0.5996 1.0000
5.000 0.7127 0.02413 0.01608 -0.0337 0.5635 1.0000
5.250 0.7200 0.02403 0.01592 -0.0286 0.5197 1.0000
5.500 0.7251 0.02406 0.01594 -0.0235 0.4772 1.0000
5.750 0.7235 0.02415 0.01588 -0.0172 0.4216 1.0000
6.000 0.7209 0.02453 0.01601 -0.0110 0.3496 1.0000
6.250 0.7139 0.02573 0.01648 -0.0048 0.2387 1.0000
6.500 0.7088 0.02775 0.01769 0.0003 0.1481 1.0000
6.750 0.7115 0.02937 0.01900 0.0044 0.1166 1.0000
7.000 0.7155 0.03083 0.02030 0.0083 0.1033 1.0000
7.250 0.7204 0.03230 0.02167 0.0120 0.0966 1.0000
7.500 0.7320 0.03372 0.02321 0.0150 0.0916 1.0000
7.750 0.7515 0.03542 0.02490 0.0168 0.0874 1.0000
8.000 0.7923 0.03794 0.02743 0.0157 0.0843 1.0000
8.250 0.8208 0.04026 0.03010 0.0164 0.0814 1.0000
8.500 0.8424 0.04278 0.03295 0.0177 0.0792 1.0000
8.750 0.8577 0.04546 0.03594 0.0198 0.0776 1.0000
9.000 0.8672 0.04844 0.03940 0.0225 0.0782 1.0000
9.250 0.8706 0.05157 0.04293 0.0258 0.0794 1.0000
9.500 0.8682 0.05483 0.04656 0.0295 0.0804 1.0000
9.750 0.8634 0.05808 0.05009 0.0330 0.0817 1.0000
10.000 0.8563 0.06141 0.05365 0.0364 0.0828 1.0000
10.250 0.8489 0.06514 0.05756 0.0396 0.0840 1.0000
10.500 0.8485 0.06912 0.06166 0.0416 0.0852 1.0000
10.750 0.8112 0.07098 0.06376 0.0474 0.0861 1.0000
11.000 0.7728 0.07400 0.06697 0.0506 0.0868 1.0000
11.250 0.7301 0.07879 0.07191 0.0507 0.0876 1.0000
11.500 0.6832 0.08607 0.07921 0.0468 0.0890 1.0000
11.750 0.6367 0.09687 0.08998 0.0394 0.0922 1.0000
12.000 0.6194 0.10508 0.09813 0.0354 0.0952 1.0000
12.250 0.6243 0.10933 0.10236 0.0355 0.0973 1.0000
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Polar data table (+)
Polar graphs
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