FX 60-100 AIRFOIL (fx60100-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: FX 60-100 AIRFOIL (fx60100-il) Reynolds number: 200,000 Max Cl/Cd: 77.73 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-fx60100-il-200000-n5.txt Download as CSV file: xf-fx60100-il-200000-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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3829 0.08107 0.07762 -0.0424 1.0000 0.0135
-8.500 -0.3901 0.07762 0.07426 -0.0425 1.0000 0.0133
-8.250 -0.3999 0.07461 0.07132 -0.0420 1.0000 0.0131
-8.000 -0.4163 0.07201 0.06881 -0.0402 1.0000 0.0129
-7.750 -0.4233 0.06663 0.06350 -0.0440 0.9959 0.0127
-7.500 -0.4056 0.05043 0.04708 -0.0696 0.9844 0.0121
-7.250 -0.3833 0.04070 0.03680 -0.0851 0.9750 0.0117
-7.000 -0.3522 0.03355 0.02897 -0.0943 0.9707 0.0113
-6.750 -0.3241 0.02900 0.02382 -0.0984 0.9648 0.0112
-6.500 -0.2910 0.02558 0.01983 -0.1017 0.9610 0.0112
-6.250 -0.2558 0.02299 0.01679 -0.1044 0.9583 0.0114
-6.000 -0.2199 0.02097 0.01442 -0.1068 0.9561 0.0118
-5.750 -0.1916 0.01956 0.01278 -0.1072 0.9500 0.0122
-5.500 -0.1583 0.01868 0.01168 -0.1085 0.9460 0.0135
-5.250 -0.1239 0.01712 0.00998 -0.1103 0.9427 0.0144
-5.000 -0.0940 0.01606 0.00885 -0.1110 0.9366 0.0147
-4.750 -0.0618 0.01516 0.00786 -0.1121 0.9316 0.0153
-4.500 -0.0275 0.01440 0.00699 -0.1136 0.9280 0.0161
-4.250 0.0020 0.01383 0.00629 -0.1141 0.9212 0.0171
-4.000 0.0346 0.01331 0.00561 -0.1150 0.9156 0.0187
-3.750 0.0658 0.01280 0.00502 -0.1156 0.9097 0.0227
-3.500 0.0964 0.01211 0.00450 -0.1164 0.9033 0.0704
-3.250 0.1288 0.01104 0.00406 -0.1181 0.8988 0.2233
-3.000 0.1569 0.01057 0.00403 -0.1185 0.8913 0.3510
-2.750 0.1871 0.01044 0.00391 -0.1188 0.8855 0.3903
-2.500 0.2154 0.01037 0.00380 -0.1187 0.8774 0.4108
-2.250 0.2450 0.01029 0.00366 -0.1187 0.8703 0.4338
-2.000 0.2728 0.01025 0.00360 -0.1185 0.8612 0.4544
-1.750 0.3018 0.01020 0.00353 -0.1184 0.8544 0.4738
-1.500 0.3296 0.01017 0.00348 -0.1182 0.8452 0.4892
-1.250 0.3577 0.01013 0.00341 -0.1180 0.8358 0.5031
-1.000 0.3861 0.01009 0.00333 -0.1178 0.8264 0.5158
-0.750 0.4137 0.01008 0.00330 -0.1175 0.8155 0.5283
-0.500 0.4416 0.01007 0.00327 -0.1173 0.8053 0.5406
-0.250 0.4695 0.01005 0.00323 -0.1170 0.7944 0.5511
0.000 0.4972 0.01004 0.00319 -0.1167 0.7822 0.5608
0.250 0.5247 0.01004 0.00318 -0.1164 0.7695 0.5722
0.500 0.5522 0.01005 0.00319 -0.1161 0.7562 0.5843
0.750 0.5795 0.01007 0.00319 -0.1158 0.7414 0.5960
1.000 0.6066 0.01010 0.00319 -0.1154 0.7240 0.6071
1.250 0.6332 0.01014 0.00321 -0.1149 0.7030 0.6186
1.500 0.6598 0.01021 0.00322 -0.1143 0.6813 0.6304
1.750 0.6860 0.01031 0.00326 -0.1138 0.6564 0.6425
2.000 0.7119 0.01043 0.00332 -0.1132 0.6299 0.6560
2.250 0.7377 0.01057 0.00343 -0.1126 0.6061 0.6715
2.500 0.7633 0.01072 0.00355 -0.1120 0.5831 0.6892
2.750 0.7886 0.01088 0.00370 -0.1114 0.5594 0.7085
3.000 0.8132 0.01107 0.00387 -0.1106 0.5335 0.7300
3.500 0.8605 0.01143 0.00422 -0.1087 0.4813 0.7945
3.750 0.8793 0.01139 0.00433 -0.1065 0.4585 0.8900
4.000 0.9055 0.01165 0.00455 -0.1062 0.4335 1.0000
4.250 0.9308 0.01201 0.00481 -0.1058 0.4083 1.0000
4.500 0.9562 0.01237 0.00509 -0.1054 0.3827 1.0000
4.750 0.9806 0.01281 0.00541 -0.1049 0.3536 1.0000
5.000 1.0049 0.01326 0.00578 -0.1044 0.3246 1.0000
5.250 1.0288 0.01376 0.00617 -0.1039 0.2955 1.0000
5.500 1.0521 0.01431 0.00660 -0.1032 0.2645 1.0000
5.750 1.0743 0.01496 0.00709 -0.1025 0.2277 1.0000
6.000 1.0959 0.01571 0.00765 -0.1017 0.1885 1.0000
6.250 1.1175 0.01644 0.00823 -0.1010 0.1585 1.0000
6.500 1.1394 0.01714 0.00882 -0.1002 0.1359 1.0000
6.750 1.1616 0.01778 0.00943 -0.0995 0.1217 1.0000
7.000 1.1840 0.01839 0.01005 -0.0988 0.1073 1.0000
7.250 1.2055 0.01907 0.01073 -0.0979 0.0955 1.0000
7.500 1.2258 0.01985 0.01147 -0.0970 0.0859 1.0000
7.750 1.2476 0.02046 0.01217 -0.0961 0.0763 1.0000
8.000 1.2681 0.02117 0.01292 -0.0951 0.0684 1.0000
8.250 1.2873 0.02199 0.01372 -0.0940 0.0623 1.0000
8.500 1.3064 0.02276 0.01461 -0.0928 0.0582 1.0000
8.750 1.3238 0.02367 0.01556 -0.0914 0.0545 1.0000
9.000 1.3381 0.02480 0.01674 -0.0897 0.0505 1.0000
9.250 1.3572 0.02547 0.01757 -0.0886 0.0471 1.0000
9.500 1.3745 0.02623 0.01843 -0.0873 0.0428 1.0000
9.750 1.3877 0.02718 0.01941 -0.0854 0.0393 1.0000
10.000 1.4038 0.02788 0.02025 -0.0839 0.0361 1.0000
10.250 1.4174 0.02874 0.02119 -0.0822 0.0333 1.0000
10.500 1.4282 0.02982 0.02228 -0.0803 0.0302 1.0000
10.750 1.4418 0.03071 0.02335 -0.0787 0.0275 1.0000
11.000 1.4536 0.03176 0.02450 -0.0770 0.0238 1.0000
11.250 1.4631 0.03303 0.02588 -0.0753 0.0206 1.0000
11.500 1.4723 0.03438 0.02733 -0.0736 0.0160 1.0000
11.750 1.4785 0.03605 0.02909 -0.0718 0.0116 1.0000
12.000 1.4816 0.03807 0.03117 -0.0700 0.0093 1.0000
12.250 1.4827 0.04035 0.03354 -0.0683 0.0079 1.0000
12.500 1.4824 0.04287 0.03618 -0.0669 0.0070 1.0000
12.750 1.4825 0.04544 0.03894 -0.0657 0.0065 1.0000
13.000 1.4809 0.04832 0.04200 -0.0648 0.0060 1.0000
13.250 1.4778 0.05151 0.04538 -0.0642 0.0057 1.0000
13.500 1.4729 0.05508 0.04914 -0.0640 0.0055 1.0000
13.750 1.4662 0.05908 0.05333 -0.0642 0.0054 1.0000
14.000 1.4576 0.06358 0.05804 -0.0650 0.0052 1.0000
14.250 1.4471 0.06862 0.06328 -0.0663 0.0051 1.0000
14.500 1.4352 0.07420 0.06909 -0.0682 0.0051 1.0000
14.750 1.4218 0.08039 0.07549 -0.0707 0.0050 1.0000
15.000 1.4067 0.08727 0.08258 -0.0738 0.0050 1.0000
15.250 1.3910 0.09465 0.09016 -0.0775 0.0050 1.0000
15.500 1.3736 0.10276 0.09847 -0.0818 0.0049 1.0000
15.750 1.3560 0.11128 0.10720 -0.0865 0.0050 1.0000
16.000 1.3376 0.12038 0.11650 -0.0918 0.0050 1.0000
16.250 1.3170 0.13041 0.12669 -0.0977 0.0051 1.0000
16.500 1.2969 0.14073 0.13719 -0.1039 0.0052 1.0000
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Polar data table (+)
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