GOE 210 (DAIMLER) AIRFOIL (goe210-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 210 (DAIMLER) AIRFOIL (goe210-il) Reynolds number: 1,000,000 Max Cl/Cd: 99.8 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe210-il-1000000-n5.txt Download as CSV file: xf-goe210-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 210 (DAIMLER) 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.3572 0.10343 0.10182 -0.0198 1.0000 0.0047
-8.750 -0.3515 0.10040 0.09880 -0.0210 1.0000 0.0043
-8.500 -0.3466 0.09707 0.09549 -0.0224 1.0000 0.0040
-8.250 -0.3429 0.09363 0.09207 -0.0236 1.0000 0.0037
-7.750 -0.3232 0.08546 0.08392 -0.0306 0.9874 0.0036
-5.750 -0.1505 0.01248 0.00800 -0.0961 0.8237 0.0054
-5.500 -0.1238 0.01158 0.00686 -0.0960 0.8112 0.0057
-5.250 -0.0967 0.01094 0.00605 -0.0960 0.8007 0.0061
-5.000 -0.0693 0.01037 0.00531 -0.0959 0.7916 0.0065
-4.500 -0.0142 0.00945 0.00410 -0.0958 0.7756 0.0075
-4.000 0.0413 0.00880 0.00329 -0.0957 0.7617 0.0099
-3.750 0.0691 0.00852 0.00290 -0.0957 0.7549 0.0109
-3.500 0.0972 0.00843 0.00282 -0.0956 0.7470 0.0132
-3.250 0.1250 0.00844 0.00278 -0.0955 0.7381 0.0151
-3.000 0.1531 0.00837 0.00268 -0.0955 0.7306 0.0166
-2.750 0.1812 0.00845 0.00275 -0.0954 0.7230 0.0180
-2.500 0.2092 0.00844 0.00271 -0.0954 0.7141 0.0196
-2.250 0.2371 0.00841 0.00262 -0.0953 0.7020 0.0215
-2.000 0.2647 0.00837 0.00248 -0.0952 0.6864 0.0228
-1.750 0.2924 0.00841 0.00246 -0.0951 0.6722 0.0237
-1.500 0.3203 0.00839 0.00238 -0.0950 0.6603 0.0241
-1.250 0.3479 0.00797 0.00187 -0.0950 0.6492 0.0259
-1.000 0.3758 0.00781 0.00166 -0.0950 0.6386 0.0269
-0.750 0.4037 0.00771 0.00150 -0.0950 0.6262 0.0275
-0.500 0.4313 0.00766 0.00136 -0.0949 0.6095 0.0279
-0.250 0.4589 0.00763 0.00125 -0.0948 0.5910 0.0281
0.000 0.4863 0.00765 0.00117 -0.0947 0.5660 0.0285
0.250 0.5134 0.00773 0.00112 -0.0946 0.5365 0.0291
0.500 0.5401 0.00787 0.00110 -0.0944 0.4974 0.0290
0.750 0.5656 0.00820 0.00115 -0.0940 0.4341 0.0291
1.000 0.5902 0.00870 0.00128 -0.0936 0.3516 0.0292
1.250 0.6159 0.00903 0.00138 -0.0933 0.2996 0.0295
1.500 0.6426 0.00922 0.00144 -0.0932 0.2701 0.0301
1.750 0.6691 0.00944 0.00153 -0.0930 0.2395 0.0313
2.000 0.6954 0.00967 0.00163 -0.0928 0.2103 0.0340
2.250 0.7219 0.00983 0.00178 -0.0926 0.1876 0.0780
2.500 0.7488 0.00987 0.00193 -0.0926 0.1738 0.1651
3.000 0.7985 0.00853 0.00233 -0.0920 0.1584 1.0000
3.250 0.8257 0.00867 0.00244 -0.0919 0.1540 1.0000
3.500 0.8526 0.00884 0.00257 -0.0918 0.1483 1.0000
3.750 0.8795 0.00901 0.00271 -0.0917 0.1432 1.0000
4.000 0.9064 0.00917 0.00285 -0.0915 0.1377 1.0000
4.250 0.9328 0.00938 0.00301 -0.0914 0.1296 1.0000
4.500 0.9591 0.00961 0.00318 -0.0912 0.1175 1.0000
4.750 0.9818 0.01033 0.00362 -0.0905 0.0646 1.0000
5.000 1.0075 0.01061 0.00386 -0.0902 0.0587 1.0000
5.250 1.0334 0.01087 0.00411 -0.0900 0.0530 1.0000
5.500 1.0595 0.01108 0.00434 -0.0897 0.0511 1.0000
5.750 1.0853 0.01133 0.00459 -0.0895 0.0481 1.0000
6.000 1.1105 0.01164 0.00488 -0.0892 0.0435 1.0000
6.250 1.1361 0.01190 0.00514 -0.0889 0.0402 1.0000
6.500 1.1590 0.01249 0.00558 -0.0882 0.0187 1.0000
6.750 1.1834 0.01289 0.00597 -0.0878 0.0118 1.0000
7.000 1.2080 0.01324 0.00638 -0.0873 0.0096 1.0000
7.250 1.2316 0.01373 0.00690 -0.0867 0.0072 1.0000
7.500 1.2558 0.01409 0.00730 -0.0862 0.0064 1.0000
7.750 1.2794 0.01453 0.00778 -0.0857 0.0056 1.0000
8.000 1.3020 0.01508 0.00838 -0.0849 0.0048 1.0000
8.250 1.3253 0.01552 0.00887 -0.0843 0.0045 1.0000
8.500 1.3481 0.01599 0.00939 -0.0836 0.0040 1.0000
8.750 1.3705 0.01649 0.00992 -0.0830 0.0036 1.0000
9.000 1.3919 0.01709 0.01057 -0.0821 0.0033 1.0000
9.250 1.4121 0.01781 0.01136 -0.0811 0.0030 1.0000
9.500 1.4329 0.01841 0.01206 -0.0801 0.0029 1.0000
9.750 1.4529 0.01907 0.01280 -0.0791 0.0027 1.0000
10.000 1.4721 0.01978 0.01358 -0.0779 0.0025 1.0000
10.250 1.4905 0.02052 0.01440 -0.0767 0.0024 1.0000
10.500 1.5081 0.02129 0.01525 -0.0754 0.0023 1.0000
10.750 1.5248 0.02208 0.01611 -0.0740 0.0021 1.0000
11.000 1.5399 0.02297 0.01708 -0.0723 0.0020 1.0000
11.250 1.5515 0.02402 0.01823 -0.0702 0.0019 1.0000
11.500 1.5555 0.02538 0.01972 -0.0669 0.0018 1.0000
11.750 1.5613 0.02661 0.02109 -0.0640 0.0018 1.0000
12.000 1.5691 0.02773 0.02231 -0.0617 0.0017 1.0000
12.250 1.5760 0.02897 0.02366 -0.0594 0.0017 1.0000
12.500 1.5808 0.03042 0.02523 -0.0572 0.0016 1.0000
12.750 1.5845 0.03204 0.02697 -0.0551 0.0016 1.0000
13.000 1.5875 0.03384 0.02889 -0.0533 0.0015 1.0000
13.250 1.5878 0.03601 0.03119 -0.0517 0.0015 1.0000
13.500 1.5875 0.03840 0.03371 -0.0504 0.0014 1.0000
13.750 1.5848 0.04118 0.03663 -0.0494 0.0014 1.0000
14.000 1.5812 0.04423 0.03982 -0.0487 0.0014 1.0000
14.250 1.5757 0.04764 0.04337 -0.0483 0.0013 1.0000
14.500 1.5684 0.05147 0.04733 -0.0483 0.0013 1.0000
14.750 1.5601 0.05560 0.05161 -0.0487 0.0013 1.0000
15.000 1.5485 0.06038 0.05655 -0.0495 0.0013 1.0000
15.250 1.5357 0.06567 0.06199 -0.0508 0.0013 1.0000
15.500 1.5206 0.07163 0.06809 -0.0527 0.0013 1.0000
15.750 1.5037 0.07829 0.07491 -0.0553 0.0012 1.0000
16.000 1.4853 0.08566 0.08242 -0.0584 0.0012 1.0000
16.250 1.4631 0.09399 0.09091 -0.0621 0.0013 1.0000
16.500 1.4409 0.10258 0.09965 -0.0660 0.0013 1.0000
16.750 1.4189 0.11125 0.10846 -0.0700 0.0013 1.0000
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Polar data table (+)
Polar graphs
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