FX 60-100 AIRFOIL (fx60100-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: FX 60-100 AIRFOIL (fx60100-il) Reynolds number: 500,000 Max Cl/Cd: 87.6 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-fx60100-il-500000-n5.txt Download as CSV file: xf-fx60100-il-500000-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 = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.3828 0.08353 0.08132 -0.0399 1.0000 0.0078
-8.750 -0.3927 0.07889 0.07673 -0.0408 1.0000 0.0074
-8.500 -0.3973 0.07269 0.07057 -0.0450 0.9977 0.0073
-8.250 -0.4185 0.04383 0.04143 -0.0779 0.9769 0.0062
-8.000 -0.3909 0.03321 0.03020 -0.0989 0.9675 0.0062
-7.750 -0.3668 0.02776 0.02422 -0.1054 0.9597 0.0061
-7.500 -0.3393 0.02398 0.01997 -0.1092 0.9543 0.0060
-7.250 -0.3131 0.02130 0.01691 -0.1110 0.9473 0.0060
-7.000 -0.2839 0.01915 0.01440 -0.1126 0.9420 0.0061
-6.750 -0.2573 0.01760 0.01257 -0.1131 0.9344 0.0062
-6.500 -0.2289 0.01636 0.01109 -0.1137 0.9281 0.0063
-6.250 -0.2023 0.01490 0.00941 -0.1140 0.9205 0.0065
-6.000 -0.1749 0.01368 0.00805 -0.1144 0.9138 0.0068
-5.750 -0.1475 0.01292 0.00721 -0.1146 0.9064 0.0071
-5.500 -0.1196 0.01229 0.00650 -0.1148 0.8996 0.0074
-5.250 -0.0916 0.01172 0.00583 -0.1149 0.8922 0.0078
-5.000 -0.0633 0.01123 0.00525 -0.1151 0.8856 0.0082
-4.750 -0.0350 0.01085 0.00481 -0.1153 0.8791 0.0089
-4.500 -0.0065 0.01048 0.00435 -0.1154 0.8726 0.0093
-4.250 0.0222 0.01008 0.00387 -0.1156 0.8663 0.0095
-4.000 0.0511 0.00967 0.00335 -0.1158 0.8594 0.0098
-3.750 0.0799 0.00936 0.00294 -0.1160 0.8526 0.0104
-3.500 0.1085 0.00913 0.00262 -0.1160 0.8445 0.0114
-3.250 0.1371 0.00894 0.00236 -0.1161 0.8367 0.0133
-3.000 0.1659 0.00860 0.00207 -0.1163 0.8288 0.0423
-2.750 0.1948 0.00821 0.00185 -0.1166 0.8207 0.1002
-2.250 0.2525 0.00725 0.00154 -0.1177 0.8022 0.3156
-2.000 0.2808 0.00716 0.00148 -0.1177 0.7923 0.3479
-1.750 0.3090 0.00714 0.00142 -0.1176 0.7819 0.3656
-1.500 0.3371 0.00711 0.00138 -0.1175 0.7692 0.3843
-1.250 0.3650 0.00711 0.00134 -0.1174 0.7544 0.4028
-1.000 0.3927 0.00713 0.00131 -0.1172 0.7348 0.4180
-0.750 0.4200 0.00718 0.00130 -0.1170 0.7134 0.4358
-0.500 0.4475 0.00723 0.00130 -0.1168 0.6930 0.4528
-0.250 0.4750 0.00731 0.00131 -0.1166 0.6724 0.4661
0.000 0.5023 0.00740 0.00134 -0.1163 0.6522 0.4791
0.250 0.5295 0.00750 0.00138 -0.1161 0.6293 0.4940
0.500 0.5562 0.00766 0.00144 -0.1158 0.6005 0.5079
0.750 0.5827 0.00785 0.00152 -0.1155 0.5699 0.5184
1.000 0.6092 0.00804 0.00160 -0.1152 0.5416 0.5282
1.250 0.6361 0.00822 0.00170 -0.1150 0.5170 0.5382
1.500 0.6631 0.00837 0.00180 -0.1148 0.4961 0.5481
1.750 0.6903 0.00852 0.00190 -0.1146 0.4786 0.5584
2.000 0.7175 0.00867 0.00203 -0.1144 0.4611 0.5697
2.250 0.7444 0.00884 0.00216 -0.1143 0.4402 0.5808
2.500 0.7709 0.00906 0.00230 -0.1140 0.4147 0.5907
2.750 0.7973 0.00929 0.00246 -0.1138 0.3908 0.6009
3.250 0.8499 0.00978 0.00284 -0.1133 0.3412 0.6238
3.500 0.8761 0.01003 0.00304 -0.1130 0.3175 0.6377
4.000 0.9277 0.01059 0.00352 -0.1124 0.2682 0.6693
4.250 0.9530 0.01092 0.00379 -0.1120 0.2398 0.6873
4.500 0.9780 0.01127 0.00409 -0.1116 0.2127 0.7086
4.750 1.0020 0.01172 0.00445 -0.1111 0.1780 0.7350
5.000 1.0260 0.01211 0.00481 -0.1105 0.1486 0.7700
5.250 1.0482 0.01235 0.00515 -0.1095 0.1305 0.8336
5.500 1.0685 0.01244 0.00537 -0.1078 0.1201 1.0000
5.750 1.0936 0.01283 0.00571 -0.1075 0.1047 1.0000
6.000 1.1182 0.01327 0.00609 -0.1070 0.0928 1.0000
6.250 1.1431 0.01367 0.00646 -0.1066 0.0849 1.0000
6.500 1.1682 0.01401 0.00682 -0.1062 0.0760 1.0000
6.750 1.1924 0.01446 0.00723 -0.1057 0.0657 1.0000
7.000 1.2164 0.01492 0.00766 -0.1052 0.0593 1.0000
7.250 1.2397 0.01543 0.00816 -0.1046 0.0532 1.0000
7.500 1.2638 0.01583 0.00860 -0.1040 0.0483 1.0000
7.750 1.2868 0.01634 0.00910 -0.1034 0.0408 1.0000
8.000 1.3098 0.01683 0.00960 -0.1027 0.0366 1.0000
8.250 1.3322 0.01735 0.01012 -0.1020 0.0328 1.0000
8.500 1.3544 0.01788 0.01069 -0.1012 0.0298 1.0000
8.750 1.3765 0.01839 0.01124 -0.1004 0.0272 1.0000
9.000 1.3976 0.01897 0.01185 -0.0995 0.0244 1.0000
9.250 1.4182 0.01958 0.01248 -0.0986 0.0214 1.0000
9.500 1.4375 0.02025 0.01315 -0.0975 0.0160 1.0000
9.750 1.4510 0.02148 0.01427 -0.0956 0.0055 1.0000
10.000 1.4668 0.02242 0.01527 -0.0940 0.0041 1.0000
10.250 1.4818 0.02335 0.01629 -0.0922 0.0035 1.0000
10.500 1.4941 0.02430 0.01736 -0.0900 0.0031 1.0000
10.750 1.5061 0.02517 0.01834 -0.0878 0.0030 1.0000
11.000 1.5170 0.02612 0.01940 -0.0856 0.0028 1.0000
11.250 1.5269 0.02717 0.02056 -0.0833 0.0027 1.0000
11.500 1.5356 0.02833 0.02185 -0.0811 0.0026 1.0000
11.750 1.5434 0.02962 0.02325 -0.0789 0.0025 1.0000
12.000 1.5502 0.03103 0.02479 -0.0769 0.0024 1.0000
12.250 1.5557 0.03261 0.02651 -0.0749 0.0023 1.0000
12.500 1.5599 0.03438 0.02841 -0.0730 0.0022 1.0000
12.750 1.5628 0.03633 0.03050 -0.0713 0.0021 1.0000
13.000 1.5637 0.03858 0.03291 -0.0697 0.0021 1.0000
13.250 1.5628 0.04112 0.03560 -0.0683 0.0020 1.0000
13.500 1.5598 0.04401 0.03865 -0.0672 0.0020 1.0000
13.750 1.5547 0.04732 0.04212 -0.0665 0.0019 1.0000
14.000 1.5472 0.05114 0.04610 -0.0662 0.0019 1.0000
14.250 1.5380 0.05544 0.05057 -0.0665 0.0019 1.0000
14.500 1.5274 0.06022 0.05553 -0.0673 0.0018 1.0000
14.750 1.5156 0.06553 0.06100 -0.0688 0.0018 1.0000
15.000 1.5047 0.07105 0.06669 -0.0707 0.0018 1.0000
15.250 1.4921 0.07720 0.07301 -0.0732 0.0018 1.0000
15.500 1.4781 0.08396 0.07994 -0.0762 0.0018 1.0000
15.750 1.4619 0.09148 0.08763 -0.0799 0.0018 1.0000
16.000 1.4445 0.09959 0.09591 -0.0840 0.0018 1.0000
16.250 1.4255 0.10837 0.10486 -0.0887 0.0018 1.0000
16.500 1.4055 0.11765 0.11432 -0.0939 0.0018 1.0000
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Polar data table (+)
Polar graphs
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