GOE 682 AIRFOIL (goe682-il) Xfoil prediction polar at RE=50,000 Ncrit=5
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Airfoil: GOE 682 AIRFOIL (goe682-il) Reynolds number: 50,000 Max Cl/Cd: 37.5 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe682-il-50000-n5.txt Download as CSV file: xf-goe682-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 682 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3347 0.10803 0.10115 -0.0363 1.0000 0.1448
-8.500 -0.3482 0.10643 0.09968 -0.0359 1.0000 0.1451
-8.250 -0.3609 0.10467 0.09805 -0.0347 1.0000 0.1452
-8.000 -0.3398 0.09521 0.08846 -0.0356 1.0000 0.0968
-7.750 -0.3914 0.09193 0.08543 -0.0387 1.0000 0.0865
-7.500 -0.3863 0.08959 0.08313 -0.0353 1.0000 0.0855
-7.250 -0.3929 0.08712 0.08074 -0.0341 1.0000 0.0849
-7.000 -0.4011 0.08449 0.07817 -0.0334 1.0000 0.0846
-6.750 -0.4091 0.08162 0.07535 -0.0334 1.0000 0.0846
-6.500 -0.4145 0.07862 0.07238 -0.0335 1.0000 0.0843
-6.250 -0.3978 0.07357 0.06721 -0.0392 0.9938 0.0839
-6.000 -0.3738 0.06820 0.06167 -0.0456 0.9856 0.0827
-5.750 -0.3494 0.06255 0.05580 -0.0522 0.9770 0.0816
-5.500 -0.3240 0.05709 0.05004 -0.0581 0.9685 0.0810
-5.250 -0.2965 0.05185 0.04436 -0.0637 0.9605 0.0828
-5.000 -0.2660 0.04686 0.03879 -0.0687 0.9533 0.0843
-4.750 -0.2391 0.04309 0.03450 -0.0714 0.9450 0.0848
-4.500 -0.2066 0.03971 0.03049 -0.0742 0.9387 0.0855
-4.250 -0.1774 0.03706 0.02723 -0.0758 0.9307 0.0865
-4.000 -0.1467 0.03552 0.02561 -0.0773 0.9239 0.0896
-3.750 -0.1164 0.03423 0.02411 -0.0785 0.9160 0.0933
-3.500 -0.0836 0.03266 0.02211 -0.0798 0.9091 0.0966
-3.250 -0.0518 0.03122 0.02026 -0.0807 0.9013 0.0994
-3.000 -0.0169 0.03005 0.01887 -0.0821 0.8948 0.1030
-2.500 0.0515 0.02848 0.01703 -0.0845 0.8797 0.1178
-2.000 0.1204 0.02723 0.01563 -0.0866 0.8640 0.1401
-1.750 0.1488 0.02678 0.01516 -0.0866 0.8540 0.1570
-1.500 0.1851 0.02619 0.01461 -0.0879 0.8471 0.1886
-1.250 0.2136 0.02569 0.01428 -0.0880 0.8376 0.2348
-1.000 0.2456 0.02513 0.01394 -0.0887 0.8299 0.3015
-0.750 0.2738 0.02457 0.01372 -0.0888 0.8210 0.3815
-0.500 0.3287 0.02247 0.01337 -0.0924 0.8173 1.0000
-0.250 0.3518 0.02272 0.01330 -0.0916 0.8059 1.0000
0.000 0.3826 0.02283 0.01311 -0.0919 0.7973 1.0000
0.500 0.4376 0.02317 0.01302 -0.0914 0.7779 1.0000
0.750 0.4716 0.02317 0.01283 -0.0920 0.7704 1.0000
1.000 0.4942 0.02348 0.01300 -0.0911 0.7591 1.0000
1.250 0.5237 0.02360 0.01297 -0.0910 0.7502 1.0000
1.500 0.5518 0.02374 0.01301 -0.0908 0.7406 1.0000
1.750 0.5758 0.02402 0.01320 -0.0900 0.7296 1.0000
2.000 0.6105 0.02395 0.01302 -0.0906 0.7222 1.0000
2.250 0.6318 0.02432 0.01334 -0.0894 0.7100 1.0000
2.500 0.6564 0.02457 0.01354 -0.0886 0.6986 1.0000
2.750 0.6868 0.02454 0.01343 -0.0884 0.6880 1.0000
3.000 0.7162 0.02444 0.01328 -0.0878 0.6753 1.0000
3.250 0.7407 0.02450 0.01329 -0.0867 0.6605 1.0000
3.500 0.7652 0.02458 0.01332 -0.0855 0.6456 1.0000
3.750 0.7892 0.02472 0.01345 -0.0844 0.6313 1.0000
4.000 0.8133 0.02492 0.01364 -0.0834 0.6177 1.0000
4.250 0.8380 0.02511 0.01382 -0.0825 0.6046 1.0000
4.500 0.8636 0.02527 0.01398 -0.0816 0.5917 1.0000
4.750 0.8897 0.02542 0.01412 -0.0809 0.5787 1.0000
5.000 0.9110 0.02576 0.01449 -0.0796 0.5642 1.0000
5.250 0.9325 0.02609 0.01487 -0.0783 0.5495 1.0000
5.500 0.9541 0.02642 0.01523 -0.0770 0.5346 1.0000
5.750 0.9757 0.02676 0.01561 -0.0757 0.5197 1.0000
6.000 0.9972 0.02711 0.01599 -0.0744 0.5046 1.0000
6.250 1.0187 0.02748 0.01641 -0.0732 0.4893 1.0000
6.500 1.0402 0.02789 0.01684 -0.0719 0.4740 1.0000
6.750 1.0614 0.02834 0.01732 -0.0706 0.4587 1.0000
7.000 1.0823 0.02886 0.01788 -0.0694 0.4436 1.0000
7.250 1.1030 0.02944 0.01850 -0.0681 0.4288 1.0000
7.500 1.1236 0.03008 0.01918 -0.0670 0.4148 1.0000
7.750 1.1444 0.03075 0.01988 -0.0658 0.4014 1.0000
8.000 1.1662 0.03141 0.02059 -0.0648 0.3889 1.0000
8.250 1.1840 0.03228 0.02158 -0.0634 0.3765 1.0000
8.500 1.2011 0.03323 0.02265 -0.0621 0.3650 1.0000
8.750 1.2209 0.03407 0.02358 -0.0609 0.3547 1.0000
9.000 1.2402 0.03493 0.02453 -0.0598 0.3446 1.0000
9.250 1.2540 0.03604 0.02585 -0.0581 0.3346 1.0000
9.500 1.2713 0.03690 0.02679 -0.0567 0.3246 1.0000
9.750 1.2856 0.03776 0.02774 -0.0549 0.3139 1.0000
10.000 1.2923 0.03893 0.02909 -0.0524 0.3034 1.0000
10.250 1.3044 0.03987 0.03012 -0.0504 0.2937 1.0000
10.500 1.3155 0.04090 0.03130 -0.0484 0.2843 1.0000
10.750 1.3214 0.04230 0.03290 -0.0461 0.2755 1.0000
11.000 1.3383 0.04313 0.03377 -0.0447 0.2663 1.0000
11.250 1.3355 0.04498 0.03589 -0.0419 0.2575 1.0000
11.500 1.3402 0.04630 0.03730 -0.0397 0.2472 1.0000
11.750 1.3404 0.04754 0.03854 -0.0372 0.2349 1.0000
12.000 1.3281 0.04985 0.04103 -0.0347 0.2235 1.0000
12.250 1.3158 0.05242 0.04371 -0.0327 0.2118 1.0000
12.500 1.3049 0.05508 0.04643 -0.0312 0.1998 1.0000
12.750 1.2933 0.05808 0.04947 -0.0303 0.1875 1.0000
13.000 1.2801 0.06175 0.05332 -0.0300 0.1758 1.0000
13.250 1.2670 0.06597 0.05777 -0.0303 0.1648 1.0000
13.500 1.2555 0.07022 0.06218 -0.0308 0.1538 1.0000
13.750 1.2443 0.07466 0.06674 -0.0317 0.1426 1.0000
14.000 1.2321 0.07945 0.07163 -0.0329 0.1308 1.0000
14.250 1.2204 0.08437 0.07658 -0.0343 0.1196 1.0000
14.500 1.2086 0.08955 0.08179 -0.0359 0.1089 1.0000
14.750 1.1961 0.09511 0.08736 -0.0378 0.0992 1.0000
15.000 1.1848 0.10054 0.09274 -0.0396 0.0908 1.0000
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Polar data table (+)
Polar graphs
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