GOE 602 MOD. AIRFOIL (goe602m-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 602 MOD. AIRFOIL (goe602m-il) Reynolds number: 50,000 Max Cl/Cd: 40.52 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe602m-il-50000-n5.txt Download as CSV file: xf-goe602m-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 602 MOD. 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.3651 0.09866 0.09193 -0.0382 1.0000 0.0586
-8.250 -0.3695 0.09575 0.08912 -0.0379 1.0000 0.0581
-8.000 -0.3764 0.09304 0.08652 -0.0372 1.0000 0.0579
-7.750 -0.3856 0.09051 0.08410 -0.0362 1.0000 0.0575
-7.500 -0.3954 0.08788 0.08160 -0.0354 1.0000 0.0572
-7.250 -0.4028 0.08473 0.07855 -0.0355 1.0000 0.0569
-7.000 -0.4088 0.08139 0.07530 -0.0362 1.0000 0.0565
-6.750 -0.4135 0.07765 0.07161 -0.0375 1.0000 0.0562
-6.500 -0.4158 0.07346 0.06746 -0.0394 1.0000 0.0557
-6.250 -0.4156 0.06872 0.06270 -0.0419 1.0000 0.0551
-6.000 -0.4123 0.06345 0.05734 -0.0449 1.0000 0.0544
-5.750 -0.4052 0.05799 0.05168 -0.0481 1.0000 0.0539
-5.500 -0.3935 0.05256 0.04593 -0.0513 1.0000 0.0537
-5.250 -0.3778 0.04791 0.04088 -0.0537 1.0000 0.0554
-5.000 -0.3582 0.04349 0.03586 -0.0560 1.0000 0.0583
-4.750 -0.3351 0.03932 0.03092 -0.0576 1.0000 0.0604
-4.500 -0.3117 0.03607 0.02706 -0.0583 1.0000 0.0620
-4.250 -0.2917 0.03457 0.02549 -0.0582 1.0000 0.0667
-4.000 -0.2627 0.03251 0.02285 -0.0592 0.9981 0.0722
-3.750 -0.2253 0.03039 0.02011 -0.0612 0.9932 0.0753
-3.500 -0.1899 0.02925 0.01878 -0.0634 0.9873 0.0832
-3.250 -0.1543 0.02805 0.01721 -0.0650 0.9810 0.0899
-3.000 -0.1181 0.02704 0.01604 -0.0668 0.9750 0.0955
-2.750 -0.0830 0.02624 0.01492 -0.0681 0.9679 0.1032
-2.500 -0.0475 0.02547 0.01401 -0.0699 0.9607 0.1164
-2.250 -0.0090 0.02455 0.01320 -0.0725 0.9541 0.1519
-2.000 0.0257 0.02364 0.01294 -0.0745 0.9461 0.2875
-1.750 0.0613 0.02313 0.01306 -0.0763 0.9390 0.4501
-1.500 0.0896 0.02283 0.01305 -0.0762 0.9290 0.5634
-1.250 0.1202 0.02236 0.01286 -0.0763 0.9205 0.6592
-1.000 0.1625 0.02150 0.01248 -0.0781 0.9121 1.0000
-0.750 0.1971 0.02167 0.01226 -0.0798 0.9014 1.0000
-0.500 0.2345 0.02182 0.01206 -0.0819 0.8919 1.0000
-0.250 0.2722 0.02194 0.01190 -0.0839 0.8825 1.0000
0.000 0.3038 0.02208 0.01182 -0.0847 0.8711 1.0000
0.250 0.3369 0.02220 0.01175 -0.0858 0.8603 1.0000
0.500 0.3747 0.02224 0.01162 -0.0876 0.8514 1.0000
0.750 0.4080 0.02232 0.01157 -0.0885 0.8405 1.0000
1.250 0.4685 0.02252 0.01159 -0.0892 0.8166 1.0000
1.500 0.5003 0.02257 0.01157 -0.0897 0.8053 1.0000
1.750 0.5358 0.02251 0.01146 -0.0907 0.7954 1.0000
2.000 0.5660 0.02256 0.01149 -0.0909 0.7830 1.0000
2.250 0.5951 0.02263 0.01155 -0.0908 0.7700 1.0000
2.500 0.6245 0.02269 0.01161 -0.0908 0.7570 1.0000
2.750 0.6543 0.02274 0.01169 -0.0908 0.7438 1.0000
3.000 0.6843 0.02279 0.01176 -0.0908 0.7305 1.0000
3.250 0.7143 0.02283 0.01184 -0.0907 0.7168 1.0000
3.500 0.7439 0.02290 0.01197 -0.0906 0.7026 1.0000
3.750 0.7729 0.02299 0.01211 -0.0903 0.6879 1.0000
4.000 0.8014 0.02311 0.01229 -0.0900 0.6727 1.0000
4.250 0.8294 0.02327 0.01255 -0.0896 0.6571 1.0000
4.500 0.8570 0.02345 0.01281 -0.0891 0.6413 1.0000
4.750 0.8843 0.02366 0.01311 -0.0886 0.6253 1.0000
5.000 0.9114 0.02391 0.01348 -0.0880 0.6093 1.0000
5.250 0.9381 0.02419 0.01387 -0.0874 0.5933 1.0000
5.500 0.9628 0.02457 0.01438 -0.0866 0.5766 1.0000
5.750 0.9862 0.02500 0.01496 -0.0856 0.5591 1.0000
6.000 1.0095 0.02539 0.01554 -0.0845 0.5406 1.0000
6.250 1.0337 0.02571 0.01596 -0.0833 0.5210 1.0000
6.500 1.0530 0.02614 0.01653 -0.0815 0.4983 1.0000
6.750 1.0733 0.02649 0.01696 -0.0797 0.4746 1.0000
7.000 1.0838 0.02685 0.01730 -0.0763 0.4381 1.0000
7.250 1.0900 0.02733 0.01762 -0.0724 0.3932 1.0000
7.500 1.1006 0.02801 0.01834 -0.0696 0.3586 1.0000
7.750 1.1098 0.02884 0.01918 -0.0667 0.3215 1.0000
8.000 1.1161 0.02986 0.02009 -0.0636 0.2775 1.0000
8.250 1.1182 0.03120 0.02119 -0.0601 0.2263 1.0000
8.500 1.1193 0.03294 0.02264 -0.0569 0.1685 1.0000
8.750 1.1140 0.03547 0.02463 -0.0537 0.1069 1.0000
9.000 1.1118 0.03803 0.02693 -0.0511 0.0726 1.0000
9.250 1.1124 0.04040 0.02924 -0.0488 0.0537 1.0000
9.500 1.1125 0.04285 0.03169 -0.0468 0.0470 1.0000
9.750 1.1112 0.04548 0.03433 -0.0450 0.0439 1.0000
10.000 1.1107 0.04811 0.03715 -0.0434 0.0417 1.0000
10.250 1.1096 0.05091 0.04013 -0.0421 0.0399 1.0000
10.500 1.1075 0.05394 0.04330 -0.0411 0.0382 1.0000
10.750 1.1045 0.05718 0.04667 -0.0404 0.0367 1.0000
11.000 1.1004 0.06066 0.05030 -0.0399 0.0354 1.0000
11.250 1.1017 0.06368 0.05357 -0.0393 0.0340 1.0000
11.500 1.1036 0.06670 0.05681 -0.0387 0.0328 1.0000
11.750 1.1074 0.06960 0.05992 -0.0380 0.0320 1.0000
12.000 1.1133 0.07236 0.06288 -0.0372 0.0312 1.0000
12.250 1.1211 0.07506 0.06581 -0.0363 0.0306 1.0000
12.500 1.1293 0.07790 0.06889 -0.0354 0.0299 1.0000
12.750 1.1348 0.08117 0.07239 -0.0350 0.0292 1.0000
13.000 1.1370 0.08488 0.07631 -0.0350 0.0286 1.0000
13.250 1.1372 0.08895 0.08055 -0.0352 0.0279 1.0000
13.500 1.1353 0.09354 0.08533 -0.0358 0.0274 1.0000
13.750 1.1266 0.09888 0.09088 -0.0376 0.0272 1.0000
14.000 1.1155 0.10467 0.09694 -0.0401 0.0272 1.0000
14.250 1.1033 0.11094 0.10344 -0.0432 0.0272 1.0000
14.500 1.0910 0.11758 0.11028 -0.0467 0.0272 1.0000
14.750 1.0783 0.12466 0.11755 -0.0507 0.0273 1.0000
15.000 1.0600 0.13373 0.12694 -0.0569 0.0279 1.0000
15.250 1.0240 0.15035 0.14390 -0.0681 0.0301 1.0000
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Polar data table (+)
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