FX 61-140 AIRFOIL (fx61140-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: FX 61-140 AIRFOIL (fx61140-il) Reynolds number: 50,000 Max Cl/Cd: 14.4 at α=1.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx61140-il-50000.txt Download as CSV file: xf-fx61140-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: FX 61-140 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.4120 0.08697 0.08094 -0.0341 1.0000 0.2177
-7.750 -0.4683 0.08017 0.07443 -0.0390 1.0000 0.1948
-7.500 -0.5202 0.07550 0.06988 -0.0429 1.0000 0.1930
-7.000 -0.5570 0.06168 0.05534 -0.0488 1.0000 0.1355
-6.750 -0.5533 0.05663 0.04987 -0.0492 1.0000 0.1210
-6.500 -0.5426 0.05244 0.04560 -0.0482 1.0000 0.1146
-6.250 -0.5315 0.04837 0.04078 -0.0488 1.0000 0.1057
-6.000 -0.5165 0.04487 0.03702 -0.0485 1.0000 0.1020
-5.750 -0.4982 0.04166 0.03334 -0.0484 1.0000 0.0992
-5.500 -0.4788 0.03914 0.03045 -0.0481 1.0000 0.0999
-5.250 -0.4585 0.03696 0.02792 -0.0476 1.0000 0.1024
-5.000 -0.4373 0.03504 0.02566 -0.0468 1.0000 0.1038
-4.750 -0.4170 0.03344 0.02378 -0.0454 1.0000 0.1054
-4.500 -0.3992 0.03218 0.02235 -0.0435 1.0000 0.1085
-4.250 -0.3857 0.03116 0.02149 -0.0411 1.0000 0.1171
-4.000 -0.3732 0.03040 0.02079 -0.0382 1.0000 0.1260
-3.750 -0.3602 0.02980 0.02013 -0.0356 1.0000 0.1363
-3.500 -0.3452 0.02877 0.01939 -0.0345 1.0000 0.1594
-3.250 -0.3241 0.02518 0.01886 -0.0350 1.0000 0.4761
-3.000 -0.3465 0.02891 0.02286 -0.0204 1.0000 0.6706
-2.750 -0.3492 0.03293 0.02670 -0.0071 0.9848 0.7435
-2.500 -0.3521 0.03504 0.02871 0.0053 0.9703 0.8032
-2.250 -0.2071 0.03683 0.02946 -0.0017 0.9693 0.9259
-2.000 -0.1237 0.03625 0.02835 -0.0139 0.9574 0.9442
-1.750 -0.0469 0.03565 0.02730 -0.0251 0.9451 0.9583
-1.500 0.0325 0.03494 0.02622 -0.0368 0.9330 0.9703
-1.250 0.1028 0.03423 0.02523 -0.0469 0.9198 0.9802
-1.000 0.1690 0.03357 0.02435 -0.0561 0.9065 0.9889
-0.750 0.2324 0.03296 0.02355 -0.0647 0.8936 0.9968
-0.500 0.2799 0.03259 0.02305 -0.0701 0.8804 1.0000
-0.250 0.3128 0.03247 0.02284 -0.0727 0.8671 1.0000
0.000 0.3451 0.03240 0.02269 -0.0750 0.8543 1.0000
0.250 0.3793 0.03233 0.02256 -0.0775 0.8422 1.0000
0.500 0.4315 0.03183 0.02197 -0.0827 0.8331 1.0000
0.750 0.4476 0.03221 0.02233 -0.0820 0.8198 1.0000
1.250 0.4637 0.03352 0.02362 -0.0782 0.7944 1.0000
1.500 0.4692 0.03428 0.02438 -0.0758 0.7828 1.0000
1.750 0.4965 0.03448 0.02458 -0.0766 0.7734 1.0000
2.000 0.5018 0.03527 0.02536 -0.0742 0.7625 1.0000
2.250 0.4757 0.03686 0.02697 -0.0675 0.7507 1.0000
2.500 0.4517 0.03828 0.02840 -0.0611 0.7406 1.0000
2.750 0.4907 0.03836 0.02848 -0.0633 0.7326 1.0000
3.000 0.4177 0.04077 0.03089 -0.0509 0.7219 1.0000
3.250 0.4069 0.04191 0.03202 -0.0469 0.7138 1.0000
3.500 0.4030 0.04295 0.03307 -0.0439 0.7056 1.0000
3.750 0.3962 0.04448 0.03460 -0.0416 0.6972 1.0000
4.000 0.4298 0.04520 0.03535 -0.0434 0.6893 1.0000
4.250 0.4254 0.04730 0.03747 -0.0426 0.6805 1.0000
4.500 0.4608 0.04828 0.03852 -0.0448 0.6727 1.0000
4.750 0.4646 0.05051 0.04079 -0.0453 0.6643 1.0000
5.000 0.4895 0.05210 0.04245 -0.0472 0.6565 1.0000
5.250 0.5053 0.05413 0.04454 -0.0486 0.6486 1.0000
5.500 0.5233 0.05621 0.04669 -0.0503 0.6411 1.0000
5.750 0.5407 0.05839 0.04899 -0.0520 0.6340 1.0000
6.000 0.5576 0.06069 0.05137 -0.0537 0.6267 1.0000
6.250 0.5720 0.06316 0.05393 -0.0554 0.6200 1.0000
6.500 0.5911 0.06554 0.05640 -0.0571 0.6129 1.0000
6.750 0.5977 0.06850 0.05946 -0.0587 0.6074 1.0000
7.000 0.6238 0.07073 0.06184 -0.0606 0.5993 1.0000
7.250 0.6236 0.07424 0.06544 -0.0621 0.5963 1.0000
7.500 0.6288 0.07776 0.06907 -0.0641 0.5954 1.0000
7.750 0.6352 0.08153 0.07296 -0.0663 0.5975 1.0000
8.000 0.6555 0.08550 0.07708 -0.0695 0.6017 1.0000
9.750 0.7947 0.10156 0.09432 -0.0747 0.4856 1.0000
10.000 0.7809 0.10688 0.09968 -0.0768 0.4839 1.0000
10.250 0.7754 0.11193 0.10487 -0.0790 0.4831 1.0000
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