FX 60-100 AIRFOIL (fx60100-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: FX 60-100 AIRFOIL (fx60100-il) Reynolds number: 100,000 Max Cl/Cd: 61.36 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-fx60100-il-100000-n5.txt Download as CSV file: xf-fx60100-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: FX 60-100 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.3672 0.09245 0.08744 -0.0393 1.0000 0.0235
-8.750 -0.3700 0.08869 0.08376 -0.0402 1.0000 0.0228
-8.500 -0.3755 0.08471 0.07987 -0.0411 1.0000 0.0222
-8.250 -0.3849 0.08048 0.07575 -0.0419 1.0000 0.0214
-7.750 -0.4292 0.06826 0.06376 -0.0448 1.0000 0.0195
-7.500 -0.4417 0.06440 0.05996 -0.0461 1.0000 0.0193
-7.250 -0.4513 0.05988 0.05544 -0.0496 1.0000 0.0191
-7.000 -0.4573 0.05449 0.04991 -0.0547 1.0000 0.0190
-6.750 -0.4398 0.04783 0.04286 -0.0635 0.9960 0.0187
-6.500 -0.4090 0.04136 0.03579 -0.0723 0.9906 0.0187
-6.250 -0.3724 0.03649 0.03002 -0.0790 0.9860 0.0196
-6.000 -0.3390 0.03246 0.02562 -0.0836 0.9825 0.0207
-5.750 -0.3061 0.02975 0.02249 -0.0863 0.9779 0.0214
-5.500 -0.2708 0.02719 0.01950 -0.0888 0.9741 0.0216
-5.250 -0.2345 0.02512 0.01710 -0.0911 0.9711 0.0220
-5.000 -0.2021 0.02351 0.01526 -0.0923 0.9666 0.0226
-4.750 -0.1685 0.02213 0.01370 -0.0937 0.9618 0.0235
-4.500 -0.1314 0.02094 0.01236 -0.0957 0.9579 0.0248
-4.250 -0.0963 0.01998 0.01123 -0.0973 0.9530 0.0265
-4.000 -0.0619 0.01910 0.01027 -0.0991 0.9476 0.0312
-3.750 -0.0233 0.01822 0.00924 -0.1016 0.9438 0.0404
-3.500 0.0156 0.01681 0.00817 -0.1045 0.9401 0.1156
-3.250 0.0489 0.01564 0.00812 -0.1069 0.9340 0.3651
-3.000 0.0865 0.01559 0.00805 -0.1085 0.9292 0.4236
-2.750 0.1180 0.01558 0.00795 -0.1090 0.9222 0.4546
-2.500 0.1518 0.01556 0.00783 -0.1100 0.9163 0.4813
-2.250 0.1852 0.01553 0.00775 -0.1108 0.9109 0.5037
-2.000 0.2153 0.01551 0.00763 -0.1110 0.9034 0.5221
-1.750 0.2515 0.01539 0.00743 -0.1123 0.8986 0.5378
-1.500 0.2795 0.01535 0.00734 -0.1121 0.8891 0.5525
-1.250 0.3129 0.01523 0.00717 -0.1129 0.8823 0.5674
-1.000 0.3425 0.01515 0.00705 -0.1129 0.8735 0.5803
-0.750 0.3729 0.01506 0.00695 -0.1131 0.8656 0.5921
-0.500 0.4038 0.01495 0.00682 -0.1134 0.8574 0.6049
-0.250 0.4322 0.01488 0.00674 -0.1131 0.8470 0.6187
0.000 0.4622 0.01475 0.00661 -0.1131 0.8373 0.6326
0.250 0.4922 0.01464 0.00650 -0.1131 0.8276 0.6459
0.500 0.5196 0.01459 0.00648 -0.1128 0.8162 0.6593
0.750 0.5478 0.01452 0.00643 -0.1124 0.8048 0.6733
1.000 0.5759 0.01442 0.00636 -0.1121 0.7924 0.6883
1.250 0.6035 0.01433 0.00632 -0.1116 0.7794 0.7049
1.500 0.6307 0.01424 0.00629 -0.1110 0.7661 0.7241
1.750 0.6573 0.01416 0.00628 -0.1103 0.7524 0.7457
2.250 0.7071 0.01394 0.00627 -0.1081 0.7207 0.8039
2.500 0.7287 0.01374 0.00622 -0.1062 0.7040 0.8629
2.750 0.7572 0.01366 0.00614 -0.1059 0.6849 1.0000
3.000 0.7853 0.01378 0.00617 -0.1057 0.6624 1.0000
3.250 0.8128 0.01393 0.00624 -0.1054 0.6376 1.0000
3.500 0.8397 0.01414 0.00635 -0.1049 0.6122 1.0000
3.750 0.8661 0.01439 0.00651 -0.1045 0.5874 1.0000
4.000 0.8922 0.01467 0.00671 -0.1040 0.5625 1.0000
4.250 0.9177 0.01500 0.00697 -0.1034 0.5359 1.0000
4.500 0.9425 0.01536 0.00726 -0.1027 0.5075 1.0000
4.750 0.9667 0.01577 0.00757 -0.1020 0.4772 1.0000
5.000 0.9906 0.01620 0.00793 -0.1012 0.4459 1.0000
5.250 1.0140 0.01668 0.00835 -0.1004 0.4143 1.0000
5.500 1.0369 0.01721 0.00880 -0.0996 0.3822 1.0000
5.750 1.0590 0.01781 0.00929 -0.0987 0.3490 1.0000
6.000 1.0808 0.01845 0.00985 -0.0978 0.3140 1.0000
6.250 1.1013 0.01921 0.01050 -0.0967 0.2793 1.0000
6.500 1.1215 0.02002 0.01120 -0.0957 0.2425 1.0000
6.750 1.1403 0.02097 0.01199 -0.0946 0.2077 1.0000
7.000 1.1587 0.02197 0.01284 -0.0934 0.1774 1.0000
7.500 1.1952 0.02401 0.01477 -0.0910 0.1325 1.0000
7.750 1.2129 0.02505 0.01581 -0.0898 0.1170 1.0000
8.000 1.2304 0.02609 0.01688 -0.0885 0.1048 1.0000
8.250 1.2466 0.02721 0.01802 -0.0870 0.0962 1.0000
8.500 1.2620 0.02838 0.01925 -0.0855 0.0890 1.0000
8.750 1.2763 0.02966 0.02059 -0.0838 0.0831 1.0000
9.000 1.2914 0.03085 0.02192 -0.0822 0.0772 1.0000
9.250 1.3041 0.03219 0.02334 -0.0804 0.0727 1.0000
9.500 1.3180 0.03338 0.02472 -0.0787 0.0681 1.0000
9.750 1.3297 0.03468 0.02611 -0.0768 0.0648 1.0000
10.000 1.3398 0.03628 0.02771 -0.0749 0.0624 1.0000
10.250 1.3539 0.03784 0.02951 -0.0734 0.0602 1.0000
10.500 1.3675 0.03952 0.03141 -0.0718 0.0579 1.0000
10.750 1.3792 0.04123 0.03329 -0.0703 0.0556 1.0000
11.000 1.3864 0.04287 0.03501 -0.0685 0.0527 1.0000
11.250 1.3913 0.04480 0.03710 -0.0668 0.0493 1.0000
11.500 1.3929 0.04644 0.03906 -0.0650 0.0450 1.0000
11.750 1.3922 0.04819 0.04096 -0.0637 0.0412 1.0000
12.000 1.3926 0.05037 0.04337 -0.0626 0.0376 1.0000
12.250 1.3942 0.05262 0.04586 -0.0619 0.0340 1.0000
12.500 1.3924 0.05524 0.04862 -0.0614 0.0314 1.0000
12.750 1.3893 0.05849 0.05206 -0.0612 0.0290 1.0000
13.000 1.3862 0.06195 0.05581 -0.0612 0.0258 1.0000
13.250 1.3801 0.06585 0.05988 -0.0619 0.0236 1.0000
13.500 1.3699 0.07056 0.06472 -0.0631 0.0223 1.0000
13.750 1.3591 0.07584 0.07024 -0.0645 0.0208 1.0000
14.000 1.3466 0.08169 0.07637 -0.0665 0.0196 1.0000
14.250 1.3325 0.08808 0.08301 -0.0691 0.0187 1.0000
14.500 1.3173 0.09510 0.09025 -0.0725 0.0181 1.0000
14.750 1.3008 0.10273 0.09810 -0.0765 0.0178 1.0000
15.000 1.2829 0.11112 0.10670 -0.0812 0.0176 1.0000
15.250 1.2634 0.12036 0.11615 -0.0866 0.0177 1.0000
15.500 1.2417 0.13073 0.12673 -0.0930 0.0181 1.0000
15.750 1.2172 0.14251 0.13871 -0.1003 0.0188 1.0000
16.000 1.1918 0.15531 0.15165 -0.1082 0.0201 1.0000
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Polar data table (+)
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