FX 66-S-171 AIRFOIL (fx66s171-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: FX 66-S-171 AIRFOIL (fx66s171-il) Reynolds number: 50,000 Max Cl/Cd: 12.66 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-fx66s171-il-50000-n5.txt Download as CSV file: xf-fx66s171-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: FX 66-S-171 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.3518 0.09493 0.08865 -0.0639 1.0000 0.0490
-11.000 -0.3629 0.08928 0.08309 -0.0665 1.0000 0.0487
-10.750 -0.3775 0.08383 0.07773 -0.0692 1.0000 0.0484
-10.500 -0.3987 0.07855 0.07256 -0.0717 1.0000 0.0481
-10.250 -0.4161 0.07306 0.06711 -0.0759 0.9811 0.0478
-10.000 -0.4258 0.06570 0.05951 -0.0855 0.9426 0.0475
-9.750 -0.4317 0.05970 0.05313 -0.0932 0.9127 0.0475
-9.500 -0.4320 0.05498 0.04794 -0.0975 0.8895 0.0478
-9.250 -0.4284 0.05100 0.04340 -0.0998 0.8704 0.0484
-9.000 -0.4172 0.04788 0.03990 -0.1008 0.8543 0.0499
-8.750 -0.4004 0.04571 0.03756 -0.1013 0.8398 0.0523
-8.500 -0.3849 0.04358 0.03511 -0.1013 0.8266 0.0549
-8.250 -0.3676 0.04133 0.03243 -0.1010 0.8152 0.0573
-8.000 -0.3484 0.03936 0.02991 -0.1004 0.8044 0.0602
-7.750 -0.3276 0.03786 0.02845 -0.0999 0.7937 0.0643
-7.500 -0.3050 0.03649 0.02683 -0.0991 0.7848 0.0693
-7.250 -0.2827 0.03522 0.02536 -0.0981 0.7758 0.0747
-7.000 -0.2637 0.03419 0.02427 -0.0971 0.7673 0.0816
-6.750 -0.2450 0.03312 0.02308 -0.0958 0.7595 0.0891
-6.500 -0.2284 0.03219 0.02200 -0.0947 0.7520 0.0999
-6.250 -0.2139 0.03111 0.02095 -0.0935 0.7449 0.1130
-6.000 -0.2000 0.02991 0.01986 -0.0926 0.7387 0.1325
-5.750 -0.1883 0.02857 0.01877 -0.0916 0.7316 0.1671
-5.500 -0.1765 0.02698 0.01790 -0.0907 0.7264 0.2549
-5.250 -0.1608 0.02711 0.01857 -0.0886 0.7195 0.3722
-5.000 -0.1385 0.02781 0.01911 -0.0870 0.7137 0.4347
-4.750 -0.1153 0.02849 0.01952 -0.0857 0.7089 0.4765
-4.500 -0.0954 0.02926 0.02016 -0.0839 0.7029 0.5065
-4.250 -0.0725 0.03002 0.02081 -0.0817 0.6982 0.5300
-4.000 -0.0483 0.03029 0.02083 -0.0806 0.6943 0.5522
-3.750 -0.0293 0.03064 0.02107 -0.0791 0.6881 0.5677
-3.500 -0.0063 0.03080 0.02108 -0.0779 0.6835 0.5801
-3.250 0.0193 0.03077 0.02083 -0.0773 0.6799 0.5913
-3.000 0.0393 0.03086 0.02077 -0.0768 0.6747 0.6016
-2.750 0.0612 0.03093 0.02070 -0.0762 0.6700 0.6103
-2.500 0.0864 0.03085 0.02042 -0.0761 0.6663 0.6187
-2.250 0.1120 0.03080 0.02016 -0.0761 0.6628 0.6269
-2.000 0.1302 0.03107 0.02037 -0.0755 0.6572 0.6338
-1.750 0.1540 0.03116 0.02031 -0.0753 0.6530 0.6415
-1.500 0.1804 0.03114 0.02012 -0.0754 0.6498 0.6490
-1.250 0.2007 0.03143 0.02032 -0.0749 0.6455 0.6563
-1.000 0.2193 0.03182 0.02064 -0.0743 0.6405 0.6637
-0.750 0.2430 0.03201 0.02072 -0.0741 0.6368 0.6714
-0.500 0.2703 0.03208 0.02066 -0.0742 0.6340 0.6797
-0.250 0.2852 0.03274 0.02131 -0.0732 0.6289 0.6877
0.000 0.3027 0.03329 0.02182 -0.0725 0.6241 0.6959
0.250 0.3270 0.03355 0.02200 -0.0724 0.6208 0.7047
0.500 0.3539 0.03369 0.02208 -0.0724 0.6182 0.7132
0.750 0.3611 0.03488 0.02329 -0.0710 0.6122 0.7221
1.000 0.3778 0.03551 0.02393 -0.0700 0.6077 0.7305
1.250 0.4034 0.03578 0.02415 -0.0701 0.6046 0.7406
1.500 0.4322 0.03588 0.02420 -0.0702 0.6024 0.7512
1.750 0.4241 0.03786 0.02632 -0.0674 0.5943 0.7610
2.000 0.4454 0.03832 0.02678 -0.0669 0.5905 0.7735
2.250 0.4739 0.03838 0.02685 -0.0669 0.5880 0.7877
2.500 0.4650 0.04047 0.02907 -0.0642 0.5798 0.8025
2.750 0.4828 0.04105 0.02973 -0.0633 0.5756 0.8207
3.000 0.5113 0.04112 0.02985 -0.0633 0.5730 0.8441
3.500 0.5338 0.04428 0.03327 -0.0636 0.5600 1.0000
3.750 0.5663 0.04473 0.03362 -0.0651 0.5575 1.0000
4.250 0.5721 0.04916 0.03793 -0.0642 0.5441 1.0000
4.500 0.6032 0.04956 0.03826 -0.0649 0.5418 1.0000
4.750 0.5840 0.05329 0.04198 -0.0635 0.5321 1.0000
5.000 0.6076 0.05418 0.04283 -0.0637 0.5288 1.0000
5.250 0.6379 0.05463 0.04325 -0.0642 0.5265 1.0000
5.500 0.6196 0.05831 0.04693 -0.0628 0.5167 1.0000
5.750 0.6430 0.05923 0.04784 -0.0629 0.5134 1.0000
6.000 0.6724 0.05974 0.04836 -0.0632 0.5112 1.0000
6.250 0.6537 0.06352 0.05215 -0.0620 0.5012 1.0000
6.500 0.6769 0.06447 0.05312 -0.0621 0.4980 1.0000
7.000 0.6865 0.06897 0.05767 -0.0613 0.4857 1.0000
7.250 0.7094 0.06999 0.05874 -0.0613 0.4826 1.0000
7.500 0.7074 0.07284 0.06163 -0.0609 0.4760 1.0000
7.750 0.7180 0.07472 0.06356 -0.0607 0.4705 1.0000
8.000 0.7414 0.07570 0.06460 -0.0607 0.4674 1.0000
8.250 0.7356 0.07895 0.06790 -0.0604 0.4602 1.0000
8.500 0.7490 0.08067 0.06971 -0.0603 0.4553 1.0000
8.750 0.7738 0.08153 0.07064 -0.0603 0.4522 1.0000
9.000 0.7645 0.08506 0.07424 -0.0601 0.4439 1.0000
9.250 0.7827 0.08642 0.07569 -0.0600 0.4395 1.0000
9.500 0.8102 0.08702 0.07641 -0.0599 0.4367 1.0000
9.750 0.7961 0.09102 0.08048 -0.0599 0.4271 1.0000
10.000 0.8189 0.09199 0.08157 -0.0598 0.4234 1.0000
10.250 0.8144 0.09531 0.08497 -0.0599 0.4154 1.0000
10.500 0.8312 0.09677 0.08655 -0.0598 0.4105 1.0000
10.750 0.8394 0.09898 0.08887 -0.0598 0.4042 1.0000
11.000 0.8470 0.10124 0.09127 -0.0599 0.3974 1.0000
11.250 0.8741 0.10163 0.09181 -0.0596 0.3938 1.0000
11.500 0.8653 0.10542 0.09569 -0.0600 0.3838 1.0000
11.750 0.8919 0.10572 0.09616 -0.0596 0.3797 1.0000
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Polar graphs
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