FX 60-100 AIRFOIL (fx60100-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: FX 60-100 AIRFOIL (fx60100-il) Reynolds number: 1,000,000 Max Cl/Cd: 96.73 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-fx60100-il-1000000-n5.txt Download as CSV file: xf-fx60100-il-1000000-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 = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.4005 0.09137 0.08971 -0.0369 1.0000 0.0039
-9.750 -0.4058 0.08633 0.08470 -0.0386 1.0000 0.0039
-9.000 -0.4760 0.03457 0.03252 -0.0939 0.9721 0.0036
-8.750 -0.4557 0.02675 0.02409 -0.1060 0.9586 0.0037
-8.500 -0.4334 0.02377 0.02076 -0.1088 0.9492 0.0038
-8.250 -0.4093 0.02179 0.01849 -0.1104 0.9412 0.0038
-8.000 -0.3845 0.02079 0.01730 -0.1109 0.9326 0.0039
-7.750 -0.3624 0.01795 0.01406 -0.1121 0.9231 0.0040
-7.500 -0.3381 0.01605 0.01185 -0.1126 0.9138 0.0040
-7.250 -0.3127 0.01454 0.01011 -0.1130 0.9048 0.0041
-7.000 -0.2866 0.01331 0.00869 -0.1133 0.8967 0.0042
-6.750 -0.2597 0.01251 0.00777 -0.1135 0.8883 0.0044
-6.500 -0.2323 0.01189 0.00703 -0.1137 0.8811 0.0045
-6.250 -0.2047 0.01134 0.00640 -0.1138 0.8738 0.0047
-6.000 -0.1768 0.01084 0.00580 -0.1140 0.8675 0.0050
-5.750 -0.1486 0.01037 0.00525 -0.1142 0.8605 0.0052
-5.500 -0.1204 0.00996 0.00476 -0.1144 0.8541 0.0055
-5.250 -0.0919 0.00958 0.00429 -0.1146 0.8474 0.0057
-5.000 -0.0635 0.00925 0.00388 -0.1148 0.8409 0.0059
-4.750 -0.0347 0.00893 0.00350 -0.1150 0.8340 0.0061
-4.500 -0.0062 0.00866 0.00316 -0.1151 0.8270 0.0062
-4.250 0.0226 0.00837 0.00280 -0.1153 0.8191 0.0065
-4.000 0.0513 0.00807 0.00241 -0.1155 0.8106 0.0070
-3.750 0.0800 0.00787 0.00214 -0.1156 0.8016 0.0072
-3.500 0.1087 0.00770 0.00191 -0.1157 0.7928 0.0075
-3.250 0.1372 0.00757 0.00172 -0.1157 0.7832 0.0080
-3.000 0.1658 0.00746 0.00154 -0.1158 0.7716 0.0088
-2.750 0.1942 0.00736 0.00138 -0.1157 0.7565 0.0124
-2.500 0.2224 0.00718 0.00121 -0.1158 0.7371 0.0445
-2.250 0.2505 0.00705 0.00109 -0.1159 0.7149 0.0774
-2.000 0.2785 0.00696 0.00100 -0.1159 0.6916 0.1124
-1.750 0.3070 0.00657 0.00088 -0.1164 0.6689 0.2255
-1.500 0.3354 0.00642 0.00084 -0.1166 0.6515 0.2861
-1.250 0.3634 0.00644 0.00082 -0.1166 0.6305 0.3098
-1.000 0.3911 0.00650 0.00083 -0.1166 0.6065 0.3294
-0.750 0.4187 0.00658 0.00086 -0.1165 0.5793 0.3539
-0.500 0.4459 0.00672 0.00089 -0.1164 0.5468 0.3748
-0.250 0.4732 0.00686 0.00094 -0.1163 0.5162 0.3885
0.000 0.5007 0.00700 0.00100 -0.1162 0.4894 0.4004
0.250 0.5285 0.00709 0.00105 -0.1161 0.4708 0.4202
0.500 0.5562 0.00720 0.00112 -0.1161 0.4513 0.4386
0.750 0.5838 0.00733 0.00119 -0.1160 0.4274 0.4516
1.000 0.6113 0.00747 0.00128 -0.1160 0.4073 0.4636
1.250 0.6389 0.00759 0.00137 -0.1159 0.3887 0.4794
1.500 0.6663 0.00775 0.00147 -0.1158 0.3669 0.4941
1.750 0.6936 0.00790 0.00158 -0.1157 0.3482 0.5050
2.000 0.7209 0.00807 0.00170 -0.1156 0.3273 0.5151
2.250 0.7474 0.00835 0.00185 -0.1154 0.2957 0.5243
2.500 0.7743 0.00856 0.00199 -0.1153 0.2740 0.5332
2.750 0.8011 0.00879 0.00215 -0.1151 0.2498 0.5436
3.000 0.8277 0.00903 0.00233 -0.1149 0.2284 0.5551
3.250 0.8545 0.00926 0.00250 -0.1147 0.2087 0.5658
3.500 0.8805 0.00958 0.00272 -0.1145 0.1813 0.5750
3.750 0.9069 0.00983 0.00292 -0.1143 0.1637 0.5846
4.000 0.9325 0.01021 0.00318 -0.1140 0.1342 0.5953
4.250 0.9590 0.01044 0.00340 -0.1138 0.1232 0.6065
4.500 0.9852 0.01070 0.00363 -0.1135 0.1075 0.6187
4.750 1.0113 0.01098 0.00388 -0.1133 0.0946 0.6319
5.000 1.0374 0.01123 0.00414 -0.1130 0.0859 0.6468
5.250 1.0642 0.01138 0.00435 -0.1129 0.0822 0.6636
5.500 1.0901 0.01166 0.00463 -0.1126 0.0709 0.6813
5.750 1.1156 0.01196 0.00494 -0.1123 0.0609 0.7013
6.250 1.1669 0.01244 0.00555 -0.1116 0.0515 0.7650
6.500 1.1922 0.01253 0.00584 -0.1112 0.0494 0.8333
6.750 1.2120 0.01253 0.00607 -0.1095 0.0453 1.0000
7.000 1.2364 0.01296 0.00647 -0.1090 0.0371 1.0000
7.250 1.2608 0.01338 0.00684 -0.1085 0.0313 1.0000
7.500 1.2856 0.01373 0.00719 -0.1081 0.0284 1.0000
7.750 1.3098 0.01414 0.00759 -0.1076 0.0253 1.0000
8.000 1.3337 0.01456 0.00800 -0.1070 0.0218 1.0000
8.250 1.3576 0.01496 0.00842 -0.1065 0.0191 1.0000
8.500 1.3802 0.01549 0.00893 -0.1058 0.0149 1.0000
8.750 1.3984 0.01651 0.00985 -0.1045 0.0037 1.0000
9.000 1.4202 0.01709 0.01048 -0.1036 0.0028 1.0000
9.250 1.4418 0.01765 0.01109 -0.1027 0.0024 1.0000
9.500 1.4627 0.01826 0.01175 -0.1017 0.0021 1.0000
9.750 1.4830 0.01890 0.01246 -0.1007 0.0019 1.0000
10.000 1.5030 0.01952 0.01315 -0.0996 0.0018 1.0000
10.250 1.5224 0.02015 0.01385 -0.0984 0.0017 1.0000
10.500 1.5409 0.02082 0.01459 -0.0971 0.0016 1.0000
10.750 1.5584 0.02154 0.01538 -0.0957 0.0016 1.0000
11.000 1.5745 0.02231 0.01622 -0.0941 0.0015 1.0000
11.250 1.5874 0.02310 0.01710 -0.0920 0.0014 1.0000
11.500 1.5981 0.02395 0.01802 -0.0895 0.0014 1.0000
11.750 1.6077 0.02488 0.01904 -0.0870 0.0013 1.0000
12.000 1.6163 0.02591 0.02016 -0.0846 0.0013 1.0000
12.500 1.6307 0.02833 0.02279 -0.0799 0.0012 1.0000
12.750 1.6368 0.02972 0.02427 -0.0777 0.0011 1.0000
13.000 1.6411 0.03131 0.02597 -0.0756 0.0011 1.0000
13.250 1.6442 0.03309 0.02786 -0.0736 0.0011 1.0000
13.500 1.6454 0.03512 0.03002 -0.0717 0.0010 1.0000
13.750 1.6447 0.03746 0.03249 -0.0701 0.0010 1.0000
14.000 1.6415 0.04017 0.03534 -0.0686 0.0010 1.0000
14.250 1.6353 0.04340 0.03871 -0.0675 0.0009 1.0000
14.500 1.6268 0.04709 0.04257 -0.0668 0.0009 1.0000
14.750 1.6174 0.05118 0.04681 -0.0666 0.0009 1.0000
15.000 1.6114 0.05509 0.05085 -0.0669 0.0009 1.0000
15.250 1.6050 0.05927 0.05517 -0.0677 0.0009 1.0000
15.500 1.5957 0.06419 0.06023 -0.0690 0.0009 1.0000
15.750 1.5859 0.06950 0.06568 -0.0707 0.0009 1.0000
16.000 1.5734 0.07558 0.07190 -0.0731 0.0009 1.0000
16.250 1.5571 0.08271 0.07920 -0.0763 0.0009 1.0000
16.500 1.5401 0.09031 0.08696 -0.0800 0.0009 1.0000
16.750 1.5194 0.09902 0.09583 -0.0845 0.0009 1.0000
17.000 1.4969 0.10850 0.10548 -0.0896 0.0009 1.0000
17.250 1.4693 0.11947 0.11662 -0.0957 0.0009 1.0000
17.500 1.4427 0.13063 0.12795 -0.1022 0.0009 1.0000
17.750 1.4178 0.14172 0.13920 -0.1087 0.0009 1.0000
|
Polar data table (+)
Polar graphs
<< Back to FX 60-100 AIRFOIL (fx60100-il)