GOE 402 AIRFOIL (goe402-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 402 AIRFOIL (goe402-il) Reynolds number: 1,000,000 Max Cl/Cd: 126.22 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe402-il-1000000-n5.txt Download as CSV file: xf-goe402-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 402 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.3015 0.11515 0.11354 -0.0197 1.0000 0.0049
-9.750 -0.2960 0.11231 0.11071 -0.0204 1.0000 0.0049
-9.500 -0.2907 0.10949 0.10790 -0.0210 1.0000 0.0049
-9.250 -0.2810 0.10617 0.10459 -0.0230 0.9992 0.0049
-9.000 -0.2729 0.10291 0.10135 -0.0245 0.9892 0.0049
-8.750 -0.2540 0.09868 0.09711 -0.0294 0.9723 0.0049
-8.500 -0.1895 0.09098 0.08928 -0.0474 0.9382 0.0045
-8.250 -0.1548 0.08624 0.08386 -0.0567 0.7850 0.0044
-8.000 -0.1502 0.08395 0.08147 -0.0570 0.7603 0.0042
-7.750 -0.1468 0.08125 0.07871 -0.0574 0.7400 0.0044
-7.500 -0.1430 0.07881 0.07624 -0.0578 0.7279 0.0044
-7.250 -0.1408 0.07653 0.07393 -0.0580 0.7170 0.0044
-7.000 -0.1337 0.07381 0.07119 -0.0597 0.7070 0.0044
-6.750 -0.1238 0.07075 0.06811 -0.0621 0.7003 0.0044
-6.250 -0.0982 0.06436 0.06166 -0.0679 0.6830 0.0045
-5.750 -0.0636 0.05713 0.05433 -0.0750 0.6673 0.0048
-5.500 -0.0433 0.05327 0.05041 -0.0789 0.6574 0.0048
-5.250 -0.0204 0.04926 0.04630 -0.0828 0.6483 0.0048
-5.000 0.0052 0.04466 0.04159 -0.0872 0.6383 0.0051
-4.750 0.0308 0.04182 0.03864 -0.0897 0.6277 0.0053
-4.500 0.0578 0.03851 0.03520 -0.0924 0.6174 0.0055
-4.250 0.0867 0.03465 0.03118 -0.0952 0.6078 0.0059
-4.000 0.1220 0.02381 0.01976 -0.0999 0.6026 0.0075
-3.750 0.1491 0.02121 0.01691 -0.1004 0.5915 0.0081
-3.500 0.1714 0.01033 0.00455 -0.1002 0.5865 0.0109
-3.000 0.2251 0.00978 0.00382 -0.0997 0.5630 0.0144
-2.750 0.2530 0.01013 0.00419 -0.0995 0.5528 0.0156
-2.500 0.2808 0.01056 0.00461 -0.0993 0.5426 0.0164
-2.250 0.3081 0.01056 0.00453 -0.0991 0.5325 0.0180
-2.000 0.3351 0.01039 0.00425 -0.0989 0.5219 0.0200
-1.750 0.3623 0.01045 0.00423 -0.0987 0.5101 0.0207
-1.500 0.3889 0.01011 0.00374 -0.0984 0.4955 0.0217
-1.250 0.4157 0.01004 0.00358 -0.0981 0.4788 0.0228
-1.000 0.4426 0.01001 0.00346 -0.0979 0.4638 0.0236
-0.750 0.4695 0.00998 0.00335 -0.0977 0.4504 0.0246
-0.500 0.4962 0.00984 0.00311 -0.0974 0.4346 0.0253
-0.250 0.5226 0.00966 0.00281 -0.0971 0.4162 0.0257
0.000 0.5487 0.00952 0.00255 -0.0967 0.3961 0.0260
0.250 0.5749 0.00946 0.00239 -0.0964 0.3760 0.0266
0.500 0.6010 0.00936 0.00219 -0.0960 0.3593 0.0268
0.750 0.6273 0.00926 0.00202 -0.0957 0.3476 0.0271
1.000 0.6537 0.00920 0.00189 -0.0954 0.3383 0.0277
1.250 0.6805 0.00917 0.00182 -0.0952 0.3312 0.0287
1.500 0.7071 0.00920 0.00181 -0.0949 0.3239 0.0297
1.750 0.7340 0.00925 0.00185 -0.0947 0.3171 0.0305
2.000 0.7607 0.00933 0.00190 -0.0945 0.3105 0.0309
2.250 0.7875 0.00929 0.00184 -0.0943 0.3063 0.0321
2.500 0.8143 0.00931 0.00183 -0.0941 0.3013 0.0332
2.750 0.8408 0.00937 0.00187 -0.0938 0.2960 0.0340
3.000 0.8677 0.00942 0.00191 -0.0936 0.2918 0.0350
3.250 0.8945 0.00947 0.00197 -0.0935 0.2875 0.0367
3.500 0.9211 0.00955 0.00205 -0.0932 0.2831 0.0386
3.750 0.9476 0.00964 0.00214 -0.0930 0.2795 0.0409
4.000 0.9744 0.00969 0.00222 -0.0929 0.2766 0.0548
4.500 1.0299 0.00843 0.00271 -0.0936 0.2692 1.0000
4.750 1.0559 0.00858 0.00284 -0.0933 0.2643 1.0000
5.000 1.0823 0.00871 0.00296 -0.0931 0.2600 1.0000
5.250 1.1083 0.00885 0.00311 -0.0928 0.2553 1.0000
5.500 1.1341 0.00903 0.00327 -0.0925 0.2498 1.0000
5.750 1.1600 0.00919 0.00343 -0.0922 0.2424 1.0000
6.000 1.1852 0.00941 0.00362 -0.0918 0.2323 1.0000
6.250 1.2072 0.00995 0.00394 -0.0910 0.1944 1.0000
6.500 1.2231 0.01112 0.00470 -0.0894 0.1240 1.0000
7.000 1.2704 0.01183 0.00537 -0.0882 0.1061 1.0000
7.250 1.2936 0.01221 0.00571 -0.0876 0.0961 1.0000
7.500 1.3106 0.01320 0.00642 -0.0861 0.0486 1.0000
7.750 1.3279 0.01415 0.00721 -0.0846 0.0187 1.0000
8.000 1.3492 0.01466 0.00772 -0.0837 0.0114 1.0000
8.250 1.3708 0.01512 0.00822 -0.0828 0.0086 1.0000
8.500 1.3918 0.01562 0.00873 -0.0819 0.0071 1.0000
8.750 1.4126 0.01610 0.00926 -0.0809 0.0060 1.0000
9.000 1.4332 0.01658 0.00978 -0.0800 0.0054 1.0000
9.250 1.4527 0.01711 0.01035 -0.0788 0.0048 1.0000
9.500 1.4704 0.01776 0.01106 -0.0774 0.0042 1.0000
9.750 1.4891 0.01830 0.01167 -0.0762 0.0039 1.0000
10.000 1.5068 0.01887 0.01229 -0.0749 0.0036 1.0000
10.250 1.5232 0.01948 0.01295 -0.0733 0.0033 1.0000
10.500 1.5370 0.02019 0.01371 -0.0714 0.0030 1.0000
10.750 1.5458 0.02099 0.01459 -0.0686 0.0028 1.0000
11.000 1.5532 0.02185 0.01553 -0.0657 0.0027 1.0000
11.250 1.5613 0.02270 0.01646 -0.0631 0.0026 1.0000
11.500 1.5684 0.02365 0.01749 -0.0606 0.0025 1.0000
11.750 1.5741 0.02475 0.01868 -0.0581 0.0024 1.0000
12.000 1.5781 0.02604 0.02007 -0.0557 0.0023 1.0000
12.250 1.5817 0.02749 0.02162 -0.0537 0.0022 1.0000
12.500 1.5842 0.02915 0.02338 -0.0520 0.0021 1.0000
12.750 1.5839 0.03126 0.02560 -0.0505 0.0021 1.0000
13.000 1.5859 0.03334 0.02778 -0.0495 0.0020 1.0000
13.250 1.5868 0.03569 0.03021 -0.0488 0.0019 1.0000
13.500 1.5851 0.03844 0.03307 -0.0483 0.0019 1.0000
13.750 1.5811 0.04161 0.03635 -0.0480 0.0018 1.0000
14.000 1.5780 0.04479 0.03963 -0.0480 0.0018 1.0000
14.250 1.5680 0.04898 0.04395 -0.0483 0.0017 1.0000
14.500 1.5584 0.05330 0.04839 -0.0489 0.0017 1.0000
14.750 1.5448 0.05835 0.05357 -0.0498 0.0016 1.0000
15.000 1.5361 0.06291 0.05825 -0.0509 0.0016 1.0000
15.250 1.5295 0.06732 0.06277 -0.0520 0.0016 1.0000
15.500 1.5173 0.07278 0.06836 -0.0536 0.0016 1.0000
15.750 1.5139 0.07697 0.07265 -0.0550 0.0015 1.0000
16.000 1.4995 0.08312 0.07893 -0.0571 0.0015 1.0000
16.250 1.4896 0.08866 0.08459 -0.0592 0.0015 1.0000
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Polar data table (+)
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