FX 60-100 (126) AIRFOIL (fx601001-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: FX 60-100 (126) AIRFOIL (fx601001-il) Reynolds number: 200,000 Max Cl/Cd: 76.76 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-fx601001-il-200000-n5.txt Download as CSV file: xf-fx601001-il-200000-n5.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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.3889 0.08645 0.08297 -0.0399 1.0000 0.0118
-9.000 -0.3939 0.08212 0.07870 -0.0412 1.0000 0.0117
-8.750 -0.4017 0.07756 0.07421 -0.0424 1.0000 0.0116
-8.500 -0.4131 0.07309 0.06983 -0.0433 1.0000 0.0115
-8.250 -0.4297 0.06877 0.06561 -0.0434 1.0000 0.0114
-8.000 -0.4525 0.06499 0.06193 -0.0421 1.0000 0.0112
-7.750 -0.4652 0.05266 0.04954 -0.0556 0.9957 0.0109
-7.500 -0.4342 0.03784 0.03392 -0.0819 0.9843 0.0108
-7.250 -0.3997 0.03138 0.02676 -0.0909 0.9780 0.0110
-7.000 -0.3656 0.02837 0.02328 -0.0951 0.9725 0.0118
-6.750 -0.3313 0.02398 0.01821 -0.0995 0.9683 0.0126
-6.500 -0.2956 0.02091 0.01467 -0.1027 0.9657 0.0131
-6.250 -0.2653 0.01892 0.01241 -0.1040 0.9602 0.0139
-6.000 -0.2312 0.01747 0.01072 -0.1058 0.9563 0.0148
-5.750 -0.1956 0.01633 0.00942 -0.1079 0.9532 0.0163
-5.500 -0.1652 0.01557 0.00854 -0.1087 0.9470 0.0190
-5.250 -0.1312 0.01472 0.00753 -0.1102 0.9421 0.0213
-5.000 -0.0955 0.01390 0.00664 -0.1121 0.9382 0.0261
-4.750 -0.0657 0.01338 0.00612 -0.1127 0.9305 0.0411
-4.500 -0.0320 0.01294 0.00564 -0.1140 0.9256 0.0581
-4.250 -0.0020 0.01258 0.00524 -0.1146 0.9180 0.0721
-4.000 0.0306 0.01212 0.00482 -0.1158 0.9119 0.0953
-3.750 0.0608 0.01161 0.00443 -0.1165 0.9036 0.1339
-3.500 0.0932 0.01095 0.00406 -0.1179 0.8969 0.2245
-3.250 0.1229 0.01050 0.00380 -0.1185 0.8882 0.2975
-3.000 0.1535 0.01025 0.00378 -0.1190 0.8807 0.3807
-2.750 0.1825 0.01026 0.00380 -0.1190 0.8712 0.4271
-2.500 0.2117 0.01033 0.00379 -0.1189 0.8617 0.4558
-2.250 0.2412 0.01040 0.00373 -0.1190 0.8527 0.4754
-2.000 0.2699 0.01048 0.00370 -0.1188 0.8427 0.4917
-1.750 0.2982 0.01055 0.00370 -0.1186 0.8323 0.5070
-1.500 0.3263 0.01056 0.00366 -0.1184 0.8220 0.5174
-1.000 0.3827 0.01053 0.00349 -0.1181 0.8002 0.5277
-0.750 0.4108 0.01055 0.00342 -0.1180 0.7887 0.5328
-0.500 0.4386 0.01055 0.00337 -0.1178 0.7770 0.5381
-0.250 0.4661 0.01057 0.00333 -0.1175 0.7629 0.5439
0.000 0.4935 0.01060 0.00328 -0.1172 0.7458 0.5499
0.250 0.5204 0.01062 0.00325 -0.1168 0.7284 0.5546
0.500 0.5478 0.01066 0.00324 -0.1166 0.7135 0.5591
0.750 0.5754 0.01072 0.00324 -0.1164 0.6995 0.5630
1.000 0.6028 0.01077 0.00325 -0.1162 0.6857 0.5658
1.250 0.6299 0.01083 0.00328 -0.1159 0.6704 0.5686
1.500 0.6568 0.01092 0.00332 -0.1155 0.6532 0.5720
1.750 0.6835 0.01104 0.00337 -0.1152 0.6342 0.5756
2.000 0.7103 0.01116 0.00344 -0.1149 0.6167 0.5788
2.250 0.7371 0.01126 0.00356 -0.1146 0.6025 0.5818
2.500 0.7639 0.01139 0.00369 -0.1143 0.5890 0.5855
2.750 0.7907 0.01152 0.00383 -0.1140 0.5753 0.5897
3.000 0.8175 0.01167 0.00400 -0.1138 0.5622 0.5938
3.250 0.8441 0.01181 0.00418 -0.1135 0.5489 0.5977
3.500 0.8707 0.01196 0.00437 -0.1132 0.5346 0.6023
4.000 0.9231 0.01230 0.00483 -0.1125 0.4993 0.6138
4.250 0.9487 0.01251 0.00506 -0.1120 0.4779 0.6209
4.500 0.9742 0.01271 0.00532 -0.1115 0.4530 0.6284
4.750 0.9986 0.01301 0.00559 -0.1109 0.4211 0.6379
5.000 1.0217 0.01341 0.00596 -0.1101 0.3834 0.6499
5.250 1.0439 0.01393 0.00637 -0.1092 0.3466 0.6658
5.500 1.0665 0.01441 0.00683 -0.1084 0.3161 0.6872
5.750 1.0878 0.01498 0.00734 -0.1074 0.2791 0.7170
6.250 1.1227 0.01592 0.00827 -0.1038 0.1991 1.0000
6.500 1.1446 0.01663 0.00885 -0.1031 0.1705 1.0000
6.750 1.1654 0.01745 0.00955 -0.1022 0.1354 1.0000
7.000 1.1825 0.01871 0.01042 -0.1010 0.0779 1.0000
7.250 1.1971 0.02026 0.01164 -0.0994 0.0320 1.0000
7.500 1.2144 0.02148 0.01278 -0.0980 0.0192 1.0000
7.750 1.2341 0.02239 0.01381 -0.0967 0.0168 1.0000
8.000 1.2526 0.02336 0.01493 -0.0954 0.0150 1.0000
8.250 1.2694 0.02444 0.01613 -0.0939 0.0132 1.0000
8.500 1.2832 0.02576 0.01759 -0.0920 0.0117 1.0000
8.750 1.2926 0.02737 0.01938 -0.0896 0.0110 1.0000
9.000 1.2963 0.02927 0.02145 -0.0864 0.0104 1.0000
9.250 1.3034 0.03078 0.02307 -0.0837 0.0102 1.0000
9.500 1.3105 0.03233 0.02475 -0.0812 0.0099 1.0000
9.750 1.3173 0.03402 0.02658 -0.0789 0.0097 1.0000
10.000 1.3243 0.03575 0.02845 -0.0767 0.0093 1.0000
10.250 1.3313 0.03751 0.03043 -0.0748 0.0088 1.0000
10.500 1.3376 0.03937 0.03243 -0.0730 0.0083 1.0000
10.750 1.3430 0.04135 0.03456 -0.0714 0.0079 1.0000
11.000 1.3470 0.04361 0.03697 -0.0699 0.0076 1.0000
11.250 1.3500 0.04611 0.03963 -0.0685 0.0075 1.0000
11.500 1.3517 0.04880 0.04249 -0.0673 0.0073 1.0000
11.750 1.3521 0.05173 0.04558 -0.0663 0.0071 1.0000
12.000 1.3509 0.05493 0.04897 -0.0656 0.0070 1.0000
12.250 1.3482 0.05844 0.05266 -0.0651 0.0069 1.0000
12.500 1.3438 0.06230 0.05671 -0.0650 0.0068 1.0000
12.750 1.3374 0.06656 0.06118 -0.0652 0.0067 1.0000
13.000 1.3290 0.07126 0.06608 -0.0659 0.0067 1.0000
13.250 1.3191 0.07638 0.07141 -0.0671 0.0066 1.0000
13.500 1.3077 0.08192 0.07718 -0.0688 0.0066 1.0000
13.750 1.2949 0.08794 0.08341 -0.0712 0.0066 1.0000
14.000 1.2819 0.09432 0.09000 -0.0741 0.0066 1.0000
14.250 1.2679 0.10121 0.09710 -0.0776 0.0066 1.0000
14.500 1.2540 0.10853 0.10462 -0.0817 0.0066 1.0000
14.750 1.2397 0.11641 0.11269 -0.0864 0.0066 1.0000
15.000 1.2256 0.12471 0.12119 -0.0917 0.0066 1.0000
15.250 1.2113 0.13367 0.13034 -0.0976 0.0067 1.0000
15.500 1.1953 0.14368 0.14054 -0.1043 0.0068 1.0000
15.750 1.1755 0.15583 0.15290 -0.1125 0.0069 1.0000
16.000 1.1390 0.17573 0.17304 -0.1252 0.0073 1.0000
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Polar data table (+)
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