GOE 289 (MVA 289) AIRFOIL (goe289-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 289 (MVA 289) AIRFOIL (goe289-il) Reynolds number: 50,000 Max Cl/Cd: 20.65 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe289-il-50000-n5.txt Download as CSV file: xf-goe289-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 289 (MVA 289) AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.250 -0.2282 0.15292 0.14619 -0.0418 1.0000 0.1219
-12.000 -0.2394 0.15240 0.14577 -0.0405 1.0000 0.1231
-11.750 -0.2450 0.15101 0.14445 -0.0420 0.9976 0.1241
-11.500 -0.2404 0.14821 0.14164 -0.0479 0.9907 0.1247
-11.250 -0.2183 0.14282 0.13625 -0.0515 0.9848 0.1257
-11.000 -0.1860 0.13732 0.13070 -0.0535 0.9790 0.1285
-10.750 -0.1680 0.13334 0.12668 -0.0574 0.9724 0.1297
-10.500 -0.1557 0.12955 0.12287 -0.0611 0.9646 0.1293
-10.000 -0.1538 0.11679 0.10999 -0.0728 0.9491 0.0852
-9.750 -0.1315 0.11271 0.10586 -0.0762 0.9438 0.0837
-9.500 -0.1234 0.10906 0.10221 -0.0789 0.9342 0.0834
-9.250 -0.1083 0.10473 0.09785 -0.0834 0.9285 0.0828
-9.000 -0.1026 0.10117 0.09429 -0.0859 0.9183 0.0827
-8.750 -0.0894 0.09669 0.08979 -0.0906 0.9125 0.0823
-8.500 -0.0860 0.09316 0.08626 -0.0928 0.9019 0.0815
-8.250 -0.0760 0.08840 0.08147 -0.0976 0.8957 0.0805
-8.000 -0.0806 0.08470 0.07779 -0.0994 0.8837 0.0797
-7.750 -0.0829 0.07973 0.07280 -0.1037 0.8750 0.0788
-7.500 -0.1003 0.07471 0.06779 -0.1064 0.8620 0.0779
-7.000 -0.1695 0.05969 0.05220 -0.1127 0.8338 0.0755
-6.750 -0.1609 0.05549 0.04771 -0.1146 0.8269 0.0759
-6.500 -0.1606 0.05339 0.04545 -0.1129 0.8155 0.0764
-6.000 -0.1335 0.04937 0.04105 -0.1124 0.7990 0.0796
-5.750 -0.1129 0.04630 0.03754 -0.1135 0.7932 0.0819
-5.500 -0.1062 0.04428 0.03512 -0.1117 0.7831 0.0834
-5.250 -0.0826 0.04143 0.03165 -0.1122 0.7772 0.0856
-5.000 -0.0646 0.04035 0.03045 -0.1111 0.7691 0.0874
-4.750 -0.0415 0.03918 0.02912 -0.1107 0.7618 0.0900
-4.500 -0.0069 0.03744 0.02700 -0.1119 0.7577 0.0951
-4.250 0.0026 0.03699 0.02641 -0.1093 0.7469 0.0989
-4.000 0.0353 0.03593 0.02521 -0.1101 0.7419 0.1065
-3.750 0.0517 0.03535 0.02449 -0.1083 0.7330 0.1132
-3.500 0.0792 0.03445 0.02338 -0.1081 0.7265 0.1252
-3.250 0.1159 0.03342 0.02229 -0.1094 0.7225 0.1436
-3.000 0.1243 0.03354 0.02229 -0.1065 0.7114 0.1574
-2.750 0.1577 0.03289 0.02171 -0.1075 0.7064 0.1832
-2.500 0.1726 0.03303 0.02186 -0.1057 0.6973 0.2066
-2.250 0.1989 0.03274 0.02156 -0.1054 0.6906 0.2308
-2.000 0.2388 0.03201 0.02076 -0.1070 0.6866 0.2530
-1.750 0.2451 0.03249 0.02123 -0.1038 0.6754 0.2664
-1.500 0.2810 0.03197 0.02065 -0.1047 0.6705 0.2912
-1.250 0.2949 0.03226 0.02094 -0.1027 0.6613 0.3134
-1.000 0.3225 0.03204 0.02072 -0.1024 0.6548 0.3464
-0.750 0.3620 0.03137 0.02001 -0.1038 0.6509 0.3938
-0.500 0.3641 0.03208 0.02079 -0.1002 0.6397 0.4226
-0.250 0.3988 0.03145 0.02027 -0.1008 0.6350 0.4751
0.000 0.4068 0.03190 0.02090 -0.0978 0.6257 0.5165
0.250 0.4325 0.03153 0.02080 -0.0970 0.6197 0.5902
0.750 0.5168 0.03111 0.02091 -0.1019 0.6056 1.0000
1.000 0.5513 0.03101 0.02047 -0.1028 0.6007 1.0000
1.250 0.5556 0.03197 0.02129 -0.0996 0.5917 1.0000
1.500 0.5808 0.03219 0.02127 -0.0991 0.5856 1.0000
1.750 0.6213 0.03190 0.02070 -0.1007 0.5820 1.0000
2.000 0.6079 0.03353 0.02230 -0.0952 0.5711 1.0000
2.250 0.6433 0.03340 0.02194 -0.0960 0.5668 1.0000
2.750 0.6633 0.03514 0.02347 -0.0914 0.5519 1.0000
3.000 0.7030 0.03482 0.02295 -0.0927 0.5484 1.0000
3.250 0.6835 0.03715 0.02529 -0.0872 0.5376 1.0000
3.500 0.7152 0.03709 0.02508 -0.0875 0.5333 1.0000
3.750 0.7570 0.03665 0.02447 -0.0889 0.5305 1.0000
4.000 0.7254 0.03989 0.02777 -0.0827 0.5183 1.0000
4.250 0.7597 0.03969 0.02744 -0.0832 0.5151 1.0000
4.750 0.7601 0.04344 0.03115 -0.0783 0.4996 1.0000
5.000 0.7966 0.04301 0.03060 -0.0787 0.4971 1.0000
5.500 0.7859 0.04810 0.03570 -0.0740 0.4810 1.0000
5.750 0.8207 0.04771 0.03521 -0.0741 0.4791 1.0000
6.250 0.7748 0.05685 0.04446 -0.0693 0.4584 1.0000
6.500 0.7618 0.06108 0.04873 -0.0681 0.4491 1.0000
6.750 0.7851 0.06155 0.04915 -0.0676 0.4466 1.0000
7.000 0.8127 0.06158 0.04914 -0.0672 0.4449 1.0000
8.000 0.7673 0.07919 0.06691 -0.0646 0.4151 1.0000
8.250 0.7882 0.08010 0.06781 -0.0642 0.4128 1.0000
8.750 0.7692 0.08874 0.07653 -0.0638 0.4002 1.0000
9.000 0.7849 0.09034 0.07813 -0.0635 0.3976 1.0000
9.250 0.8041 0.09157 0.07938 -0.0632 0.3957 1.0000
9.500 0.7754 0.09821 0.08610 -0.0635 0.3880 1.0000
9.750 0.7822 0.10076 0.08869 -0.0634 0.3840 1.0000
10.000 0.7968 0.10252 0.09047 -0.0632 0.3814 1.0000
10.250 0.8147 0.10400 0.09197 -0.0631 0.3795 1.0000
10.500 0.7975 0.10945 0.09750 -0.0636 0.3743 1.0000
10.750 0.7961 0.11296 0.10107 -0.0639 0.3696 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 289 (MVA 289) AIRFOIL (goe289-il)