FX 60-100 AIRFOIL (fx60100-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: FX 60-100 AIRFOIL (fx60100-il) Reynolds number: 1,000,000 Max Cl/Cd: 112.4 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx60100-il-1000000.txt Download as CSV file: xf-fx60100-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.2842 0.10478 0.10320 -0.0359 1.0000 0.0080
-10.750 -0.2848 0.10054 0.09897 -0.0368 1.0000 0.0087
-10.500 -0.2853 0.09574 0.09419 -0.0380 1.0000 0.0093
-5.750 -0.1402 0.01350 0.00894 -0.1165 0.9189 0.0098
-5.500 -0.1133 0.01189 0.00721 -0.1164 0.9128 0.0089
-5.250 -0.0856 0.01088 0.00609 -0.1165 0.9060 0.0090
-5.000 -0.0576 0.01030 0.00541 -0.1165 0.8998 0.0093
-4.750 -0.0293 0.00969 0.00471 -0.1166 0.8933 0.0096
-4.500 -0.0007 0.00887 0.00377 -0.1169 0.8867 0.0100
-4.250 0.0281 0.00839 0.00321 -0.1172 0.8799 0.0104
-4.000 0.0567 0.00806 0.00279 -0.1173 0.8726 0.0110
-3.750 0.0854 0.00779 0.00247 -0.1174 0.8657 0.0116
-3.500 0.1142 0.00756 0.00216 -0.1174 0.8583 0.0120
-3.250 0.1430 0.00738 0.00191 -0.1175 0.8512 0.0125
-3.000 0.1716 0.00723 0.00168 -0.1174 0.8431 0.0137
-2.750 0.2009 0.00672 0.00138 -0.1178 0.8338 0.0825
-2.500 0.2298 0.00617 0.00117 -0.1183 0.8234 0.2016
-2.250 0.2583 0.00584 0.00105 -0.1186 0.8122 0.2877
-2.000 0.2868 0.00569 0.00099 -0.1187 0.8007 0.3336
-1.750 0.3153 0.00556 0.00098 -0.1188 0.7898 0.3869
-1.500 0.3436 0.00554 0.00095 -0.1188 0.7800 0.4083
-1.250 0.3719 0.00555 0.00093 -0.1188 0.7684 0.4219
-0.750 0.4280 0.00559 0.00090 -0.1186 0.7380 0.4473
-0.500 0.4558 0.00564 0.00090 -0.1184 0.7172 0.4604
-0.250 0.4833 0.00572 0.00091 -0.1182 0.6938 0.4715
0.000 0.5109 0.00582 0.00093 -0.1181 0.6714 0.4823
0.250 0.5385 0.00591 0.00097 -0.1179 0.6510 0.4968
0.500 0.5660 0.00601 0.00102 -0.1178 0.6278 0.5113
0.750 0.5934 0.00613 0.00108 -0.1176 0.6048 0.5230
1.000 0.6208 0.00626 0.00115 -0.1175 0.5833 0.5352
1.250 0.6483 0.00638 0.00123 -0.1174 0.5627 0.5493
1.500 0.6758 0.00650 0.00132 -0.1173 0.5431 0.5626
1.750 0.7032 0.00664 0.00140 -0.1171 0.5244 0.5741
2.000 0.7306 0.00678 0.00151 -0.1170 0.5024 0.5857
2.250 0.7576 0.00696 0.00161 -0.1168 0.4785 0.5971
2.500 0.7848 0.00712 0.00173 -0.1167 0.4562 0.6084
2.750 0.8119 0.00730 0.00185 -0.1165 0.4329 0.6208
3.000 0.8387 0.00750 0.00200 -0.1164 0.4091 0.6341
3.250 0.8655 0.00770 0.00215 -0.1162 0.3837 0.6479
3.500 0.8918 0.00795 0.00232 -0.1159 0.3548 0.6625
3.750 0.9180 0.00821 0.00251 -0.1157 0.3269 0.6795
4.000 0.9434 0.00857 0.00275 -0.1154 0.2886 0.6996
4.250 0.9691 0.00886 0.00299 -0.1151 0.2573 0.7256
4.500 0.9942 0.00920 0.00326 -0.1147 0.2238 0.7627
4.750 1.0180 0.00947 0.00357 -0.1140 0.1889 0.8336
5.000 1.0373 0.00960 0.00380 -0.1122 0.1632 1.0000
5.250 1.0624 0.01004 0.00410 -0.1118 0.1365 1.0000
5.500 1.0883 0.01037 0.00437 -0.1116 0.1240 1.0000
5.750 1.1141 0.01070 0.00465 -0.1113 0.1096 1.0000
6.000 1.1393 0.01109 0.00496 -0.1109 0.0952 1.0000
6.250 1.1646 0.01146 0.00528 -0.1105 0.0856 1.0000
6.500 1.1900 0.01180 0.00559 -0.1102 0.0743 1.0000
6.750 1.2145 0.01224 0.00594 -0.1097 0.0618 1.0000
7.000 1.2391 0.01264 0.00632 -0.1093 0.0562 1.0000
7.250 1.2638 0.01304 0.00670 -0.1088 0.0506 1.0000
7.500 1.2894 0.01329 0.00701 -0.1085 0.0477 1.0000
7.750 1.3131 0.01377 0.00745 -0.1079 0.0397 1.0000
8.000 1.3371 0.01418 0.00784 -0.1074 0.0351 1.0000
8.250 1.3604 0.01466 0.00828 -0.1068 0.0298 1.0000
8.500 1.3843 0.01505 0.00870 -0.1062 0.0274 1.0000
8.750 1.4068 0.01558 0.00920 -0.1055 0.0231 1.0000
9.000 1.4295 0.01604 0.00966 -0.1048 0.0181 1.0000
9.250 1.4456 0.01723 0.01070 -0.1032 0.0054 1.0000
9.500 1.4661 0.01791 0.01144 -0.1021 0.0042 1.0000
9.750 1.4856 0.01863 0.01223 -0.1009 0.0037 1.0000
10.000 1.5039 0.01946 0.01315 -0.0995 0.0032 1.0000
10.250 1.5218 0.02027 0.01406 -0.0981 0.0030 1.0000
10.500 1.5393 0.02106 0.01494 -0.0966 0.0029 1.0000
10.750 1.5552 0.02193 0.01590 -0.0949 0.0028 1.0000
11.000 1.5690 0.02285 0.01691 -0.0930 0.0027 1.0000
11.250 1.5790 0.02380 0.01796 -0.0904 0.0026 1.0000
11.500 1.5869 0.02484 0.01910 -0.0876 0.0025 1.0000
11.750 1.5938 0.02599 0.02035 -0.0848 0.0025 1.0000
12.000 1.5998 0.02725 0.02172 -0.0822 0.0024 1.0000
12.250 1.6045 0.02868 0.02325 -0.0796 0.0023 1.0000
12.500 1.6081 0.03026 0.02495 -0.0772 0.0023 1.0000
12.750 1.6108 0.03200 0.02680 -0.0750 0.0022 1.0000
13.000 1.6112 0.03403 0.02896 -0.0729 0.0022 1.0000
13.250 1.6097 0.03636 0.03142 -0.0710 0.0021 1.0000
13.500 1.6065 0.03901 0.03420 -0.0693 0.0021 1.0000
13.750 1.6015 0.04201 0.03734 -0.0681 0.0020 1.0000
14.000 1.5949 0.04540 0.04088 -0.0672 0.0020 1.0000
14.250 1.5870 0.04919 0.04481 -0.0668 0.0020 1.0000
14.500 1.5778 0.05343 0.04920 -0.0670 0.0020 1.0000
14.750 1.5673 0.05815 0.05407 -0.0677 0.0020 1.0000
15.000 1.5552 0.06346 0.05953 -0.0691 0.0020 1.0000
15.250 1.5422 0.06931 0.06553 -0.0710 0.0020 1.0000
15.500 1.5286 0.07562 0.07200 -0.0735 0.0020 1.0000
15.750 1.5142 0.08243 0.07896 -0.0766 0.0020 1.0000
16.000 1.4986 0.08981 0.08649 -0.0801 0.0020 1.0000
16.250 1.4819 0.09771 0.09454 -0.0841 0.0020 1.0000
16.500 1.4651 0.10594 0.10292 -0.0884 0.0020 1.0000
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Polar data table (+)
Polar graphs
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