GOE 210 (DAIMLER) AIRFOIL (goe210-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
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Airfoil: GOE 210 (DAIMLER) AIRFOIL (goe210-il) Reynolds number: 1,000,000 Max Cl/Cd: 116.69 at α=1.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe210-il-1000000.txt Download as CSV file: xf-goe210-il-1000000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3667 0.08778 0.08625 -0.0227 1.0000 0.0097
-7.500 -0.3680 0.08501 0.08352 -0.0233 1.0000 0.0097
-7.250 -0.3683 0.08211 0.08064 -0.0242 1.0000 0.0097
-7.000 -0.3588 0.07704 0.07558 -0.0278 0.9986 0.0100
-6.750 -0.3338 0.07325 0.07178 -0.0336 0.9962 0.0103
-6.500 -0.3049 0.06903 0.06754 -0.0409 0.9938 0.0107
-6.250 -0.2752 0.06457 0.06304 -0.0483 0.9899 0.0113
-6.000 -0.2393 0.05938 0.05781 -0.0573 0.9864 0.0125
-5.750 -0.1921 0.05292 0.05123 -0.0688 0.9816 0.0132
-5.500 -0.1607 0.04796 0.04616 -0.0744 0.9735 0.0132
-5.250 -0.1303 0.04304 0.04112 -0.0790 0.9643 0.0133
-5.000 -0.1029 0.03534 0.03316 -0.0840 0.9521 0.0136
-4.750 -0.0807 0.03371 0.03145 -0.0847 0.9368 0.0141
-4.500 -0.0572 0.03214 0.02977 -0.0851 0.9159 0.0150
-4.250 -0.0232 0.01439 0.01050 -0.0880 0.8939 0.0140
-4.000 0.0024 0.01496 0.01102 -0.0875 0.8672 0.0145
-3.750 0.0285 0.01423 0.01006 -0.0871 0.8456 0.0154
-3.000 0.1097 0.01081 0.00577 -0.0863 0.8020 0.0195
-2.750 0.1371 0.01082 0.00575 -0.0863 0.7909 0.0208
-2.500 0.1648 0.01040 0.00519 -0.0861 0.7799 0.0222
-2.250 0.1924 0.01024 0.00489 -0.0858 0.7667 0.0235
-2.000 0.2195 0.00926 0.00373 -0.0856 0.7542 0.0254
-1.750 0.2469 0.00895 0.00339 -0.0855 0.7433 0.0271
-1.500 0.2747 0.00861 0.00300 -0.0854 0.7348 0.0282
-1.000 0.3304 0.00811 0.00240 -0.0852 0.7184 0.0311
-0.750 0.3582 0.00792 0.00216 -0.0851 0.7095 0.0319
-0.500 0.3861 0.00775 0.00194 -0.0850 0.6997 0.0323
-0.250 0.4140 0.00759 0.00175 -0.0849 0.6897 0.0326
0.000 0.4419 0.00733 0.00141 -0.0849 0.6799 0.0337
0.250 0.4698 0.00717 0.00118 -0.0848 0.6672 0.0356
0.500 0.4976 0.00710 0.00106 -0.0847 0.6511 0.0378
0.750 0.5252 0.00709 0.00099 -0.0845 0.6332 0.0398
1.000 0.5529 0.00710 0.00094 -0.0844 0.6163 0.0417
1.250 0.5804 0.00715 0.00092 -0.0843 0.5954 0.0452
1.500 0.6075 0.00714 0.00096 -0.0842 0.5679 0.1057
1.750 0.6301 0.00540 0.00118 -0.0837 0.5323 1.0000
2.000 0.6559 0.00576 0.00127 -0.0833 0.4758 1.0000
2.250 0.6807 0.00627 0.00143 -0.0829 0.4026 1.0000
2.500 0.7061 0.00670 0.00161 -0.0826 0.3470 1.0000
2.750 0.7317 0.00710 0.00177 -0.0823 0.3001 1.0000
3.000 0.7575 0.00748 0.00194 -0.0820 0.2583 1.0000
3.250 0.7834 0.00783 0.00211 -0.0817 0.2227 1.0000
3.500 0.8095 0.00814 0.00229 -0.0815 0.1982 1.0000
3.750 0.8360 0.00838 0.00245 -0.0813 0.1842 1.0000
4.000 0.8628 0.00858 0.00260 -0.0812 0.1757 1.0000
4.250 0.8896 0.00877 0.00276 -0.0810 0.1684 1.0000
4.500 0.9163 0.00896 0.00293 -0.0809 0.1619 1.0000
4.750 0.9429 0.00916 0.00310 -0.0807 0.1547 1.0000
5.000 0.9691 0.00941 0.00329 -0.0805 0.1420 1.0000
5.250 0.9952 0.00967 0.00347 -0.0803 0.1256 1.0000
5.750 1.0427 0.01080 0.00424 -0.0792 0.0626 1.0000
6.000 1.0681 0.01114 0.00456 -0.0788 0.0574 1.0000
6.250 1.0939 0.01139 0.00483 -0.0786 0.0550 1.0000
6.500 1.1199 0.01161 0.00509 -0.0783 0.0532 1.0000
6.750 1.1453 0.01191 0.00539 -0.0780 0.0496 1.0000
7.000 1.1700 0.01228 0.00574 -0.0776 0.0436 1.0000
7.250 1.1948 0.01262 0.00599 -0.0773 0.0310 1.0000
7.500 1.2163 0.01343 0.00669 -0.0763 0.0136 1.0000
7.750 1.2404 0.01387 0.00721 -0.0758 0.0124 1.0000
8.000 1.2635 0.01443 0.00784 -0.0751 0.0110 1.0000
8.250 1.2847 0.01526 0.00878 -0.0740 0.0095 1.0000
8.500 1.3073 0.01583 0.00941 -0.0733 0.0091 1.0000
8.750 1.3297 0.01641 0.01006 -0.0725 0.0086 1.0000
9.000 1.3513 0.01706 0.01079 -0.0717 0.0082 1.0000
9.250 1.3721 0.01777 0.01158 -0.0707 0.0077 1.0000
9.500 1.3920 0.01855 0.01243 -0.0696 0.0073 1.0000
9.750 1.4100 0.01950 0.01346 -0.0683 0.0069 1.0000
10.000 1.4158 0.02168 0.01584 -0.0653 0.0063 1.0000
10.250 1.4364 0.02220 0.01643 -0.0643 0.0061 1.0000
10.500 1.4534 0.02304 0.01735 -0.0629 0.0059 1.0000
10.750 1.4679 0.02404 0.01845 -0.0612 0.0057 1.0000
11.000 1.4801 0.02517 0.01968 -0.0592 0.0055 1.0000
11.250 1.4886 0.02633 0.02094 -0.0566 0.0053 1.0000
11.500 1.4946 0.02758 0.02229 -0.0538 0.0052 1.0000
11.750 1.4994 0.02894 0.02376 -0.0511 0.0050 1.0000
12.000 1.5031 0.03046 0.02540 -0.0485 0.0049 1.0000
12.250 1.5062 0.03211 0.02715 -0.0463 0.0048 1.0000
12.500 1.5085 0.03394 0.02909 -0.0443 0.0047 1.0000
12.750 1.5098 0.03598 0.03124 -0.0426 0.0046 1.0000
13.000 1.5094 0.03833 0.03371 -0.0413 0.0045 1.0000
13.250 1.5073 0.04104 0.03653 -0.0403 0.0044 1.0000
13.500 1.5030 0.04413 0.03974 -0.0396 0.0044 1.0000
13.750 1.4956 0.04778 0.04353 -0.0391 0.0043 1.0000
14.000 1.4852 0.05195 0.04783 -0.0391 0.0042 1.0000
14.250 1.4724 0.05665 0.05268 -0.0394 0.0042 1.0000
14.500 1.4570 0.06190 0.05809 -0.0401 0.0041 1.0000
14.750 1.4389 0.06783 0.06418 -0.0414 0.0041 1.0000
15.000 1.4210 0.07405 0.07057 -0.0433 0.0041 1.0000
15.250 1.3990 0.08133 0.07803 -0.0458 0.0040 1.0000
15.500 1.3856 0.08770 0.08453 -0.0487 0.0040 1.0000
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Polar data table (+)
Polar graphs
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