GOE 210 (DAIMLER) AIRFOIL (goe210-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 210 (DAIMLER) AIRFOIL (goe210-il) Reynolds number: 500,000 Max Cl/Cd: 105.13 at α=2.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe210-il-500000.txt Download as CSV file: xf-goe210-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.250 -0.3590 0.08450 0.08240 -0.0217 1.0000 0.0150
-7.000 -0.3571 0.08185 0.07979 -0.0226 1.0000 0.0155
-6.750 -0.3558 0.07923 0.07720 -0.0233 1.0000 0.0160
-6.500 -0.3537 0.07657 0.07458 -0.0245 1.0000 0.0168
-6.250 -0.3138 0.07096 0.06890 -0.0387 0.9969 0.0177
-6.000 -0.2750 0.06529 0.06315 -0.0484 0.9937 0.0178
-5.750 -0.2381 0.05977 0.05754 -0.0564 0.9893 0.0178
-5.500 -0.2104 0.05277 0.05048 -0.0633 0.9857 0.0183
-5.250 -0.1780 0.04914 0.04679 -0.0681 0.9834 0.0188
-5.000 -0.1449 0.04550 0.04307 -0.0728 0.9777 0.0196
-4.750 -0.1057 0.04138 0.03881 -0.0785 0.9732 0.0212
-4.500 -0.0554 0.03734 0.03444 -0.0838 0.9681 0.0238
-4.250 -0.0214 0.03348 0.03031 -0.0863 0.9589 0.0240
-4.000 0.0038 0.02759 0.02419 -0.0892 0.9476 0.0253
-3.750 0.0308 0.02564 0.02210 -0.0900 0.9349 0.0263
-3.500 0.0583 0.02368 0.01995 -0.0904 0.9202 0.0282
-3.250 0.0901 0.02317 0.01910 -0.0895 0.9045 0.0325
-3.000 0.1160 0.01815 0.01355 -0.0901 0.8890 0.0344
-2.750 0.1412 0.01750 0.01285 -0.0900 0.8728 0.0374
-2.500 0.1700 0.01840 0.01343 -0.0890 0.8565 0.0439
-2.250 0.1982 0.01284 0.00717 -0.0887 0.8441 0.0350
-2.000 0.2249 0.01245 0.00671 -0.0885 0.8309 0.0385
-1.750 0.2525 0.01127 0.00526 -0.0880 0.8187 0.0380
-1.500 0.2794 0.01057 0.00439 -0.0874 0.8044 0.0385
-1.250 0.3062 0.01017 0.00386 -0.0870 0.7899 0.0407
-1.000 0.3334 0.00971 0.00331 -0.0866 0.7783 0.0416
-0.750 0.3607 0.00937 0.00288 -0.0863 0.7695 0.0423
-0.500 0.3882 0.00907 0.00254 -0.0861 0.7596 0.0430
-0.250 0.4157 0.00887 0.00229 -0.0859 0.7496 0.0442
0.000 0.4432 0.00859 0.00193 -0.0857 0.7399 0.0461
0.250 0.4709 0.00843 0.00174 -0.0855 0.7295 0.0490
0.500 0.4987 0.00834 0.00164 -0.0853 0.7203 0.0527
0.750 0.5263 0.00829 0.00157 -0.0852 0.7107 0.0609
1.000 0.5527 0.00726 0.00166 -0.0855 0.6994 0.5149
1.250 0.5774 0.00618 0.00165 -0.0845 0.6848 1.0000
1.500 0.6045 0.00627 0.00164 -0.0842 0.6680 1.0000
1.750 0.6316 0.00635 0.00164 -0.0839 0.6504 1.0000
2.000 0.6587 0.00646 0.00166 -0.0836 0.6322 1.0000
2.250 0.6855 0.00659 0.00168 -0.0833 0.6103 1.0000
2.500 0.7117 0.00677 0.00173 -0.0829 0.5766 1.0000
2.750 0.7369 0.00708 0.00181 -0.0824 0.5271 1.0000
3.000 0.7611 0.00753 0.00195 -0.0818 0.4642 1.0000
3.250 0.7849 0.00809 0.00218 -0.0811 0.3979 1.0000
3.500 0.8090 0.00863 0.00245 -0.0806 0.3416 1.0000
3.750 0.8336 0.00912 0.00271 -0.0802 0.2952 1.0000
4.000 0.8583 0.00958 0.00296 -0.0798 0.2554 1.0000
4.250 0.8834 0.00999 0.00321 -0.0794 0.2266 1.0000
4.500 0.9090 0.01031 0.00346 -0.0791 0.2092 1.0000
4.750 0.9347 0.01062 0.00371 -0.0788 0.1977 1.0000
5.000 0.9606 0.01089 0.00396 -0.0785 0.1882 1.0000
5.250 0.9864 0.01116 0.00422 -0.0783 0.1800 1.0000
5.500 1.0118 0.01148 0.00451 -0.0779 0.1698 1.0000
5.750 1.0374 0.01177 0.00477 -0.0776 0.1589 1.0000
6.000 1.0629 0.01204 0.00503 -0.0773 0.1464 1.0000
6.250 1.0881 0.01237 0.00531 -0.0770 0.1302 1.0000
6.500 1.1103 0.01307 0.00574 -0.0763 0.0830 1.0000
6.750 1.1330 0.01370 0.00627 -0.0756 0.0689 1.0000
7.000 1.1573 0.01410 0.00671 -0.0751 0.0643 1.0000
7.250 1.1807 0.01460 0.00721 -0.0746 0.0590 1.0000
7.500 1.2050 0.01498 0.00764 -0.0741 0.0534 1.0000
7.750 1.2281 0.01550 0.00814 -0.0735 0.0439 1.0000
8.000 1.2500 0.01616 0.00864 -0.0727 0.0222 1.0000
8.250 1.2709 0.01697 0.00948 -0.0716 0.0171 1.0000
8.500 1.2914 0.01782 0.01042 -0.0705 0.0151 1.0000
8.750 1.3107 0.01880 0.01157 -0.0692 0.0135 1.0000
9.000 1.3305 0.01963 0.01252 -0.0680 0.0129 1.0000
9.250 1.3486 0.02061 0.01362 -0.0667 0.0123 1.0000
9.500 1.3653 0.02169 0.01482 -0.0651 0.0117 1.0000
9.750 1.3800 0.02289 0.01613 -0.0633 0.0112 1.0000
10.000 1.3926 0.02422 0.01758 -0.0613 0.0108 1.0000
10.250 1.4021 0.02575 0.01923 -0.0589 0.0104 1.0000
10.500 1.4079 0.02747 0.02107 -0.0561 0.0101 1.0000
10.750 1.4070 0.02960 0.02332 -0.0524 0.0098 1.0000
11.000 1.4025 0.03249 0.02638 -0.0486 0.0094 1.0000
11.250 1.4101 0.03377 0.02781 -0.0465 0.0092 1.0000
11.500 1.4168 0.03520 0.02940 -0.0445 0.0089 1.0000
11.750 1.4200 0.03715 0.03150 -0.0424 0.0088 1.0000
12.000 1.4216 0.03937 0.03389 -0.0404 0.0087 1.0000
12.250 1.4217 0.04183 0.03651 -0.0387 0.0086 1.0000
12.500 1.4203 0.04456 0.03942 -0.0373 0.0085 1.0000
12.750 1.4171 0.04759 0.04263 -0.0362 0.0084 1.0000
13.000 1.4122 0.05094 0.04619 -0.0355 0.0083 1.0000
13.250 1.4053 0.05466 0.05010 -0.0351 0.0082 1.0000
13.500 1.3967 0.05875 0.05438 -0.0352 0.0082 1.0000
13.750 1.3857 0.06334 0.05917 -0.0358 0.0082 1.0000
14.000 1.3735 0.06835 0.06437 -0.0369 0.0082 1.0000
14.250 1.3593 0.07389 0.07009 -0.0386 0.0082 1.0000
14.500 1.3438 0.07999 0.07639 -0.0410 0.0082 1.0000
14.750 1.3275 0.08666 0.08323 -0.0441 0.0082 1.0000
15.000 1.3102 0.09397 0.09073 -0.0478 0.0082 1.0000
15.250 1.2927 0.10183 0.09875 -0.0522 0.0083 1.0000
15.500 1.2744 0.11022 0.10730 -0.0569 0.0084 1.0000
15.750 1.2566 0.11887 0.11610 -0.0620 0.0084 1.0000
16.000 1.2379 0.12803 0.12540 -0.0674 0.0085 1.0000
16.250 1.2195 0.13745 0.13495 -0.0730 0.0086 1.0000
16.500 1.2008 0.14727 0.14487 -0.0789 0.0087 1.0000
16.750 1.1818 0.15767 0.15538 -0.0852 0.0089 1.0000
17.000 1.1608 0.16920 0.16700 -0.0921 0.0091 1.0000
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Polar data table (+)
Polar graphs
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