FX 60-100 (126) AIRFOIL (fx601001-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: FX 60-100 (126) AIRFOIL (fx601001-il) Reynolds number: 500,000 Max Cl/Cd: 103 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx601001-il-500000.txt Download as CSV file: xf-fx601001-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: FX 60-100 (126) AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.3076 0.08919 0.08705 -0.0403 1.0000 0.0242
-9.750 -0.3100 0.08466 0.08255 -0.0414 1.0000 0.0243
-7.250 -0.3931 0.02018 0.01558 -0.0992 0.9831 0.0137
-7.000 -0.3620 0.01813 0.01327 -0.1009 0.9791 0.0140
-6.750 -0.3279 0.01662 0.01158 -0.1028 0.9760 0.0146
-6.500 -0.2924 0.01535 0.01015 -0.1049 0.9738 0.0152
-6.250 -0.2557 0.01422 0.00887 -0.1070 0.9720 0.0160
-6.000 -0.2179 0.01349 0.00805 -0.1093 0.9706 0.0173
-5.750 -0.1823 0.01236 0.00679 -0.1113 0.9679 0.0183
-5.500 -0.1502 0.01138 0.00573 -0.1125 0.9627 0.0201
-5.250 -0.1153 0.01080 0.00508 -0.1141 0.9591 0.0225
-5.000 -0.0798 0.01015 0.00437 -0.1157 0.9560 0.0275
-4.750 -0.0500 0.00971 0.00399 -0.1161 0.9494 0.0452
-4.500 -0.0187 0.00938 0.00369 -0.1168 0.9437 0.0615
-4.250 0.0111 0.00917 0.00342 -0.1171 0.9373 0.0697
-4.000 0.0400 0.00885 0.00310 -0.1173 0.9295 0.0820
-3.750 0.0689 0.00845 0.00279 -0.1175 0.9218 0.1078
-3.500 0.0979 0.00759 0.00241 -0.1184 0.9138 0.2470
-3.250 0.1261 0.00701 0.00222 -0.1187 0.9048 0.3677
-3.000 0.1544 0.00693 0.00223 -0.1186 0.8961 0.4263
-2.750 0.1823 0.00698 0.00223 -0.1183 0.8858 0.4503
-2.500 0.2103 0.00706 0.00222 -0.1180 0.8756 0.4669
-2.250 0.2383 0.00714 0.00222 -0.1177 0.8655 0.4786
-2.000 0.2662 0.00721 0.00221 -0.1174 0.8549 0.4879
-1.750 0.2941 0.00730 0.00220 -0.1172 0.8432 0.4971
-1.500 0.3216 0.00733 0.00220 -0.1169 0.8309 0.5045
-1.250 0.3491 0.00740 0.00219 -0.1165 0.8168 0.5131
-1.000 0.3763 0.00747 0.00220 -0.1161 0.8012 0.5217
-0.750 0.4036 0.00751 0.00218 -0.1158 0.7856 0.5287
-0.500 0.4312 0.00754 0.00215 -0.1155 0.7710 0.5327
-0.250 0.4589 0.00758 0.00212 -0.1153 0.7566 0.5364
0.000 0.4865 0.00763 0.00209 -0.1151 0.7408 0.5404
0.250 0.5137 0.00766 0.00207 -0.1148 0.7236 0.5444
0.500 0.5411 0.00771 0.00207 -0.1146 0.7049 0.5481
0.750 0.5684 0.00780 0.00207 -0.1143 0.6858 0.5519
1.000 0.5956 0.00787 0.00208 -0.1141 0.6687 0.5557
1.250 0.6230 0.00794 0.00212 -0.1139 0.6542 0.5595
1.500 0.6506 0.00802 0.00217 -0.1138 0.6409 0.5630
1.750 0.6782 0.00811 0.00223 -0.1137 0.6280 0.5660
2.000 0.7057 0.00821 0.00229 -0.1135 0.6148 0.5688
2.250 0.7331 0.00828 0.00236 -0.1134 0.6021 0.5714
2.500 0.7605 0.00837 0.00244 -0.1132 0.5901 0.5741
2.750 0.7879 0.00848 0.00255 -0.1131 0.5783 0.5771
3.000 0.8152 0.00860 0.00266 -0.1129 0.5660 0.5805
3.250 0.8426 0.00870 0.00277 -0.1128 0.5532 0.5836
3.500 0.8698 0.00879 0.00290 -0.1126 0.5396 0.5868
3.750 0.8968 0.00892 0.00305 -0.1124 0.5244 0.5903
4.000 0.9234 0.00907 0.00319 -0.1121 0.5034 0.5943
4.250 0.9496 0.00926 0.00334 -0.1118 0.4771 0.5987
4.500 0.9754 0.00947 0.00353 -0.1114 0.4426 0.6034
4.750 0.9989 0.00996 0.00378 -0.1107 0.3866 0.6087
5.000 1.0215 0.01058 0.00415 -0.1100 0.3311 0.6144
5.250 1.0452 0.01112 0.00453 -0.1094 0.2892 0.6218
5.500 1.0695 0.01157 0.00490 -0.1089 0.2549 0.6314
5.750 1.0936 0.01202 0.00528 -0.1084 0.2236 0.6450
6.000 1.1176 0.01249 0.00569 -0.1079 0.1948 0.6657
6.250 1.1413 0.01294 0.00614 -0.1074 0.1677 0.7005
6.500 1.1638 0.01340 0.00663 -0.1067 0.1348 0.7660
6.750 1.1750 0.01427 0.00737 -0.1038 0.0634 1.0000
7.000 1.1919 0.01572 0.00851 -0.1023 0.0230 1.0000
7.250 1.2145 0.01642 0.00926 -0.1014 0.0193 1.0000
7.500 1.2358 0.01723 0.01010 -0.1004 0.0167 1.0000
7.750 1.2546 0.01831 0.01128 -0.0990 0.0151 1.0000
8.000 1.2752 0.01909 0.01215 -0.0978 0.0144 1.0000
8.250 1.2943 0.01999 0.01316 -0.0965 0.0136 1.0000
8.500 1.3120 0.02099 0.01424 -0.0950 0.0130 1.0000
8.750 1.3282 0.02207 0.01540 -0.0934 0.0124 1.0000
9.000 1.3423 0.02331 0.01670 -0.0915 0.0118 1.0000
9.250 1.3505 0.02512 0.01861 -0.0888 0.0111 1.0000
9.500 1.3571 0.02721 0.02083 -0.0860 0.0106 1.0000
9.750 1.3700 0.02820 0.02193 -0.0839 0.0104 1.0000
10.000 1.3800 0.02957 0.02341 -0.0815 0.0102 1.0000
10.250 1.3892 0.03107 0.02505 -0.0792 0.0099 1.0000
10.500 1.3978 0.03274 0.02685 -0.0769 0.0097 1.0000
10.750 1.4058 0.03456 0.02881 -0.0747 0.0095 1.0000
11.000 1.4127 0.03653 0.03093 -0.0726 0.0093 1.0000
11.250 1.4183 0.03869 0.03325 -0.0706 0.0091 1.0000
11.500 1.4221 0.04107 0.03580 -0.0686 0.0090 1.0000
11.750 1.4238 0.04369 0.03861 -0.0668 0.0089 1.0000
12.000 1.4228 0.04668 0.04181 -0.0650 0.0088 1.0000
12.250 1.4189 0.05002 0.04537 -0.0635 0.0088 1.0000
12.500 1.4105 0.05405 0.04967 -0.0623 0.0088 1.0000
12.750 1.3984 0.05866 0.05454 -0.0616 0.0089 1.0000
13.000 1.3827 0.06393 0.06008 -0.0616 0.0090 1.0000
13.250 1.3642 0.06990 0.06632 -0.0624 0.0092 1.0000
13.500 1.3450 0.07631 0.07296 -0.0642 0.0094 1.0000
13.750 1.3247 0.08328 0.08015 -0.0669 0.0095 1.0000
14.000 1.3045 0.09068 0.08776 -0.0704 0.0096 1.0000
14.250 1.2831 0.09889 0.09616 -0.0748 0.0098 1.0000
14.500 1.2629 0.10748 0.10492 -0.0800 0.0099 1.0000
14.750 1.2415 0.11708 0.11469 -0.0861 0.0100 1.0000
15.000 1.2206 0.12737 0.12515 -0.0930 0.0101 1.0000
15.250 1.1993 0.13853 0.13645 -0.1007 0.0102 1.0000
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Polar data table (+)
Polar graphs
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