GOE 546 AIRFOIL (goe546-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 546 AIRFOIL (goe546-il) Reynolds number: 200,000 Max Cl/Cd: 80.47 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe546-il-200000.txt Download as CSV file: xf-goe546-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 546 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.3049 0.08817 0.08485 -0.0385 1.0000 0.0613
-9.000 -0.3180 0.08549 0.08223 -0.0378 1.0000 0.0630
-8.750 -0.3413 0.08309 0.07992 -0.0374 1.0000 0.0642
-8.500 -0.3676 0.08084 0.07776 -0.0369 1.0000 0.0647
-8.250 -0.3914 0.07852 0.07551 -0.0359 1.0000 0.0649
-8.000 -0.4185 0.07629 0.07335 -0.0346 1.0000 0.0649
-7.750 -0.4290 0.06655 0.06343 -0.0498 0.9933 0.0652
-7.500 -0.4020 0.06510 0.06217 -0.0385 0.9940 0.0673
-7.250 -0.4870 0.06515 0.06168 -0.0523 0.9943 0.0660
-7.000 -0.4653 0.06329 0.05999 -0.0505 0.9918 0.0675
-6.750 -0.4418 0.06172 0.05843 -0.0510 0.9881 0.0704
-6.500 -0.4220 0.05156 0.04750 -0.0653 0.9795 0.0796
-6.250 -0.3966 0.04880 0.04492 -0.0661 0.9761 0.0818
-6.000 -0.3660 0.04585 0.04187 -0.0694 0.9728 0.0866
-5.750 -0.3452 0.04115 0.03670 -0.0728 0.9658 0.0957
-5.500 -0.3150 0.03107 0.02517 -0.0756 0.9613 0.0671
-5.250 -0.2790 0.02768 0.02145 -0.0777 0.9591 0.0616
-5.000 -0.2545 0.02562 0.01903 -0.0774 0.9524 0.0620
-4.750 -0.2197 0.02374 0.01676 -0.0788 0.9486 0.0628
-4.500 -0.1816 0.02206 0.01475 -0.0808 0.9461 0.0634
-4.250 -0.1522 0.02140 0.01384 -0.0810 0.9400 0.0650
-4.000 -0.1187 0.01956 0.01185 -0.0822 0.9359 0.0668
-3.750 -0.0803 0.01837 0.01062 -0.0843 0.9330 0.0690
-3.500 -0.0389 0.01751 0.00975 -0.0869 0.9309 0.0726
-3.250 -0.0122 0.01699 0.00918 -0.0865 0.9232 0.0761
-3.000 0.0268 0.01615 0.00831 -0.0886 0.9197 0.0794
-2.750 0.0689 0.01520 0.00745 -0.0914 0.9174 0.0863
-2.500 0.1023 0.01457 0.00685 -0.0924 0.9113 0.0968
-2.250 0.1373 0.01387 0.00623 -0.0936 0.9053 0.1198
-2.000 0.1709 0.01218 0.00584 -0.0953 0.9014 0.4315
-1.750 0.1962 0.01188 0.00579 -0.0944 0.8925 0.5220
-1.500 0.2304 0.01148 0.00558 -0.0953 0.8872 0.5902
-1.250 0.2565 0.01119 0.00551 -0.0945 0.8790 0.6547
-1.000 0.2831 0.01069 0.00542 -0.0934 0.8720 0.7608
-0.750 0.3644 0.01010 0.00518 -0.1032 0.8712 0.9779
-0.500 0.4064 0.01006 0.00500 -0.1062 0.8628 1.0000
-0.250 0.4358 0.01000 0.00481 -0.1063 0.8545 1.0000
0.000 0.4579 0.01005 0.00478 -0.1050 0.8430 1.0000
0.250 0.4841 0.01006 0.00470 -0.1044 0.8335 1.0000
0.500 0.5108 0.01006 0.00461 -0.1039 0.8239 1.0000
0.750 0.5341 0.01012 0.00461 -0.1028 0.8127 1.0000
1.000 0.5603 0.01013 0.00454 -0.1022 0.8026 1.0000
1.250 0.5872 0.01011 0.00444 -0.1017 0.7921 1.0000
1.500 0.6109 0.01013 0.00442 -0.1005 0.7791 1.0000
1.750 0.6355 0.01015 0.00440 -0.0996 0.7664 1.0000
2.000 0.6607 0.01017 0.00436 -0.0987 0.7536 1.0000
2.250 0.6860 0.01020 0.00433 -0.0979 0.7403 1.0000
2.500 0.7118 0.01025 0.00434 -0.0972 0.7277 1.0000
2.750 0.7376 0.01033 0.00438 -0.0965 0.7158 1.0000
3.000 0.7622 0.01044 0.00445 -0.0957 0.7020 1.0000
3.250 0.7870 0.01053 0.00451 -0.0947 0.6865 1.0000
3.500 0.8113 0.01065 0.00454 -0.0937 0.6687 1.0000
3.750 0.8342 0.01079 0.00464 -0.0924 0.6488 1.0000
4.000 0.8573 0.01096 0.00475 -0.0912 0.6288 1.0000
4.250 0.8804 0.01117 0.00488 -0.0900 0.6087 1.0000
4.500 0.9028 0.01137 0.00506 -0.0887 0.5883 1.0000
4.750 0.9249 0.01156 0.00524 -0.0875 0.5674 1.0000
5.000 0.9466 0.01178 0.00543 -0.0861 0.5451 1.0000
5.250 0.9664 0.01201 0.00559 -0.0843 0.5151 1.0000
5.500 0.9850 0.01231 0.00579 -0.0824 0.4788 1.0000
5.750 1.0034 0.01269 0.00605 -0.0805 0.4440 1.0000
6.000 1.0201 0.01319 0.00640 -0.0783 0.4039 1.0000
6.250 1.0351 0.01379 0.00680 -0.0760 0.3539 1.0000
6.500 1.0450 0.01471 0.00734 -0.0729 0.2826 1.0000
6.750 1.0491 0.01611 0.00819 -0.0691 0.1950 1.0000
7.000 1.0481 0.01805 0.00939 -0.0647 0.0970 1.0000
7.250 1.0588 0.01915 0.01040 -0.0619 0.0795 1.0000
7.500 1.0696 0.02009 0.01132 -0.0591 0.0727 1.0000
7.750 1.0823 0.02085 0.01215 -0.0566 0.0684 1.0000
8.000 1.0923 0.02179 0.01308 -0.0538 0.0648 1.0000
8.250 1.1009 0.02289 0.01419 -0.0509 0.0621 1.0000
8.500 1.1142 0.02376 0.01515 -0.0487 0.0600 1.0000
8.750 1.1268 0.02475 0.01620 -0.0465 0.0579 1.0000
9.000 1.1398 0.02578 0.01726 -0.0445 0.0560 1.0000
9.250 1.1523 0.02701 0.01847 -0.0426 0.0537 1.0000
9.500 1.1697 0.02864 0.02010 -0.0414 0.0519 1.0000
9.750 1.1887 0.02979 0.02136 -0.0402 0.0507 1.0000
10.000 1.2079 0.03100 0.02268 -0.0392 0.0489 1.0000
10.250 1.2262 0.03225 0.02402 -0.0381 0.0471 1.0000
10.500 1.2459 0.03367 0.02547 -0.0373 0.0453 1.0000
10.750 1.2933 0.03749 0.02934 -0.0410 0.0429 1.0000
11.000 1.3008 0.03861 0.03070 -0.0382 0.0421 1.0000
11.250 1.3107 0.04019 0.03253 -0.0360 0.0410 1.0000
11.500 1.3223 0.04228 0.03487 -0.0342 0.0400 1.0000
11.750 1.3315 0.04464 0.03750 -0.0323 0.0393 1.0000
12.000 1.3369 0.04697 0.04007 -0.0301 0.0385 1.0000
12.250 1.3418 0.04920 0.04247 -0.0281 0.0376 1.0000
12.500 1.3490 0.05127 0.04460 -0.0266 0.0364 1.0000
12.750 1.3524 0.05457 0.04805 -0.0252 0.0356 1.0000
13.000 1.3470 0.05905 0.05279 -0.0233 0.0352 1.0000
13.250 1.3346 0.06291 0.05692 -0.0209 0.0351 1.0000
13.500 1.3200 0.06658 0.06084 -0.0189 0.0351 1.0000
13.750 1.3037 0.07029 0.06479 -0.0173 0.0351 1.0000
14.000 1.2863 0.07412 0.06886 -0.0163 0.0352 1.0000
14.250 1.2671 0.07818 0.07315 -0.0160 0.0354 1.0000
14.500 1.2431 0.08297 0.07820 -0.0166 0.0356 1.0000
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Polar data table (+)
Polar graphs
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