GOE 382 AIRFOIL (goe382-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 382 AIRFOIL (goe382-il) Reynolds number: 200,000 Max Cl/Cd: 60.87 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe382-il-200000.txt Download as CSV file: xf-goe382-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 382 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.0333 0.10221 0.09835 -0.0867 0.8408 0.0774
-10.500 -0.0289 0.09951 0.09557 -0.0869 0.8298 0.0789
-10.250 -0.0693 0.09109 0.08716 -0.0973 0.8196 0.0831
-10.000 -0.0486 0.08985 0.08584 -0.0942 0.8092 0.0837
-9.750 -0.0328 0.08827 0.08420 -0.0928 0.7983 0.0846
-9.500 -0.0225 0.08610 0.08197 -0.0928 0.7887 0.0861
-9.250 -0.0194 0.08296 0.07879 -0.0944 0.7797 0.0883
-9.000 -0.0562 0.07351 0.06933 -0.1050 0.7735 0.0927
-8.750 -0.0390 0.07236 0.06815 -0.1033 0.7644 0.0935
-8.500 -0.0287 0.07029 0.06601 -0.1032 0.7574 0.0949
-8.250 -0.1112 0.05795 0.05335 -0.1173 0.7510 0.1021
-8.000 -0.0859 0.05647 0.05193 -0.1165 0.7438 0.1030
-7.750 -0.0684 0.05502 0.05042 -0.1158 0.7380 0.1044
-7.500 -0.1444 0.03616 0.02965 -0.1151 0.7350 0.0663
-7.250 -0.1271 0.03388 0.02723 -0.1144 0.7283 0.0651
-7.000 -0.1105 0.03177 0.02477 -0.1133 0.7228 0.0644
-6.750 -0.0917 0.02995 0.02258 -0.1123 0.7178 0.0639
-6.500 -0.0719 0.02833 0.02068 -0.1113 0.7116 0.0632
-6.250 -0.0497 0.02687 0.01890 -0.1105 0.7061 0.0626
-6.000 -0.0252 0.02569 0.01743 -0.1099 0.7014 0.0625
-5.750 -0.0009 0.02475 0.01633 -0.1094 0.6959 0.0627
-5.500 0.0242 0.02390 0.01535 -0.1089 0.6901 0.0631
-5.250 0.0507 0.02313 0.01441 -0.1086 0.6852 0.0636
-5.000 0.0779 0.02249 0.01360 -0.1084 0.6806 0.0644
-4.750 0.1028 0.02199 0.01303 -0.1078 0.6747 0.0657
-4.500 0.1293 0.02163 0.01248 -0.1074 0.6692 0.0671
-4.250 0.1569 0.02080 0.01168 -0.1074 0.6647 0.0685
-4.000 0.1830 0.02033 0.01122 -0.1071 0.6598 0.0701
-3.750 0.2080 0.01990 0.01081 -0.1066 0.6539 0.0719
-3.500 0.2346 0.01954 0.01037 -0.1062 0.6487 0.0740
-3.250 0.2607 0.01897 0.00981 -0.1059 0.6443 0.0768
-3.000 0.2838 0.01862 0.00953 -0.1052 0.6389 0.0805
-2.750 0.3067 0.01813 0.00910 -0.1044 0.6332 0.0863
-2.500 0.3309 0.01759 0.00858 -0.1038 0.6283 0.0981
-2.250 0.3559 0.01709 0.00836 -0.1034 0.6238 0.1657
-1.750 0.4001 0.01637 0.00830 -0.1018 0.6124 0.3149
-1.500 0.4243 0.01596 0.00818 -0.1013 0.6077 0.4060
-1.250 0.4436 0.01557 0.00836 -0.0997 0.6027 0.5615
-1.000 0.4649 0.01546 0.00849 -0.0983 0.5965 0.6417
-0.750 0.4891 0.01534 0.00852 -0.0973 0.5914 0.7029
-0.500 0.5155 0.01532 0.00857 -0.0965 0.5871 0.7600
-0.250 0.5368 0.01539 0.00882 -0.0948 0.5811 0.8137
0.000 0.5627 0.01546 0.00896 -0.0938 0.5752 0.8641
0.250 0.5962 0.01556 0.00897 -0.0943 0.5702 0.9002
0.500 0.6308 0.01579 0.00913 -0.0954 0.5649 0.9278
0.750 0.6678 0.01598 0.00929 -0.0971 0.5585 0.9497
1.000 0.7140 0.01611 0.00928 -0.1007 0.5530 0.9660
1.250 0.7662 0.01630 0.00930 -0.1056 0.5475 0.9771
1.500 0.8179 0.01645 0.00943 -0.1108 0.5403 0.9888
1.750 0.8779 0.01650 0.00933 -0.1175 0.5346 0.9992
2.000 0.9009 0.01661 0.00929 -0.1171 0.5303 1.0000
2.250 0.9119 0.01677 0.00950 -0.1145 0.5245 1.0000
2.500 0.9287 0.01687 0.00953 -0.1128 0.5194 1.0000
2.750 0.9498 0.01695 0.00948 -0.1117 0.5151 1.0000
3.000 0.9665 0.01716 0.00966 -0.1100 0.5102 1.0000
3.250 0.9817 0.01736 0.00986 -0.1079 0.5049 1.0000
3.500 1.0020 0.01750 0.00992 -0.1067 0.5004 1.0000
3.750 1.0269 0.01766 0.00993 -0.1063 0.4965 1.0000
4.000 1.0415 0.01795 0.01025 -0.1042 0.4915 1.0000
4.250 1.0588 0.01817 0.01046 -0.1025 0.4864 1.0000
4.500 1.0814 0.01833 0.01053 -0.1017 0.4820 1.0000
4.750 1.1095 0.01856 0.01062 -0.1020 0.4782 1.0000
5.000 1.1223 0.01892 0.01106 -0.0996 0.4734 1.0000
5.250 1.1418 0.01920 0.01134 -0.0984 0.4688 1.0000
5.500 1.1660 0.01943 0.01150 -0.0980 0.4648 1.0000
5.750 1.1965 0.01970 0.01162 -0.0988 0.4611 1.0000
6.000 1.2083 0.02012 0.01215 -0.0964 0.4565 1.0000
6.250 1.2269 0.02046 0.01251 -0.0951 0.4520 1.0000
6.500 1.2509 0.02074 0.01273 -0.0948 0.4482 1.0000
6.750 1.2808 0.02104 0.01293 -0.0955 0.4448 1.0000
7.000 1.2971 0.02155 0.01350 -0.0940 0.4410 1.0000
7.250 1.3118 0.02202 0.01404 -0.0922 0.4369 1.0000
7.500 1.3325 0.02239 0.01440 -0.0914 0.4330 1.0000
7.750 1.3595 0.02267 0.01463 -0.0917 0.4294 1.0000
8.000 1.3889 0.02312 0.01500 -0.0925 0.4260 1.0000
8.250 1.3950 0.02375 0.01577 -0.0894 0.4223 1.0000
8.500 1.4074 0.02429 0.01637 -0.0874 0.4187 1.0000
8.750 1.4274 0.02472 0.01682 -0.0866 0.4153 1.0000
9.000 1.4563 0.02505 0.01708 -0.0873 0.4122 1.0000
9.250 1.4880 0.02555 0.01752 -0.0885 0.4089 1.0000
9.500 1.4836 0.02639 0.01854 -0.0840 0.4054 1.0000
9.750 1.4913 0.02716 0.01940 -0.0816 0.4019 1.0000
10.000 1.5085 0.02773 0.02001 -0.0806 0.3987 1.0000
10.250 1.5356 0.02809 0.02033 -0.0810 0.3958 1.0000
10.500 1.5768 0.02835 0.02050 -0.0836 0.3928 1.0000
10.750 1.5673 0.02958 0.02192 -0.0789 0.3896 1.0000
11.000 1.5635 0.03082 0.02331 -0.0754 0.3861 1.0000
11.250 1.5730 0.03171 0.02427 -0.0737 0.3828 1.0000
11.500 1.5978 0.03197 0.02451 -0.0738 0.3793 1.0000
11.750 1.6460 0.03168 0.02405 -0.0768 0.3754 1.0000
12.000 1.6155 0.03382 0.02645 -0.0705 0.3717 1.0000
12.250 1.6063 0.03543 0.02820 -0.0671 0.3672 1.0000
12.500 1.6272 0.03559 0.02833 -0.0667 0.3627 1.0000
12.750 1.6747 0.03489 0.02746 -0.0691 0.3587 1.0000
13.000 1.6380 0.03799 0.03085 -0.0635 0.3549 1.0000
13.250 1.6257 0.04030 0.03331 -0.0607 0.3508 1.0000
13.500 1.6371 0.04128 0.03434 -0.0598 0.3474 1.0000
13.750 1.6679 0.04110 0.03411 -0.0604 0.3445 1.0000
14.000 1.6870 0.04174 0.03476 -0.0602 0.3412 1.0000
14.250 1.6374 0.04689 0.04022 -0.0556 0.3368 1.0000
14.500 1.6280 0.04953 0.04297 -0.0539 0.3324 1.0000
14.750 1.6624 0.04861 0.04198 -0.0544 0.3287 1.0000
15.000 1.6776 0.04933 0.04269 -0.0540 0.3246 1.0000
15.250 1.6102 0.05748 0.05118 -0.0506 0.3193 1.0000
15.500 1.6201 0.05865 0.05237 -0.0503 0.3147 1.0000
15.750 1.6760 0.05538 0.04893 -0.0512 0.3107 1.0000
16.000 1.5999 0.06542 0.05932 -0.0488 0.3051 1.0000
16.250 1.5803 0.07001 0.06403 -0.0483 0.2995 1.0000
16.500 1.6364 0.06620 0.06009 -0.0488 0.2959 1.0000
16.750 1.4268 0.09446 0.08888 -0.0480 0.2826 1.0000
17.000 1.5474 0.08166 0.07593 -0.0478 0.2832 1.0000
17.250 1.6108 0.07660 0.07072 -0.0480 0.2801 1.0000
17.500 1.4609 0.09842 0.09295 -0.0487 0.2680 1.0000
17.750 1.5408 0.09067 0.08509 -0.0484 0.2667 1.0000
18.000 1.6092 0.08471 0.07897 -0.0482 0.2637 1.0000
18.250 1.4944 0.10260 0.09721 -0.0497 0.2524 1.0000
18.500 1.5576 0.09694 0.09144 -0.0493 0.2497 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 382 AIRFOIL (goe382-il)