GOE 458 AIRFOIL (goe458-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 458 AIRFOIL (goe458-il) Reynolds number: 1,000,000 Max Cl/Cd: 140.9 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe458-il-1000000.txt Download as CSV file: xf-goe458-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 458 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.2472 0.11804 0.11652 -0.0348 1.0000 0.0049
-11.000 -0.2463 0.11546 0.11395 -0.0345 1.0000 0.0050
-10.750 -0.2464 0.11303 0.11155 -0.0339 1.0000 0.0051
-10.500 -0.2415 0.10971 0.10823 -0.0349 0.9996 0.0051
-10.250 -0.2327 0.10584 0.10436 -0.0372 0.9988 0.0053
-10.000 -0.2239 0.10187 0.10039 -0.0394 0.9979 0.0055
-9.750 -0.2151 0.09787 0.09639 -0.0417 0.9967 0.0057
-9.500 -0.2061 0.09379 0.09231 -0.0442 0.9952 0.0060
-9.250 -0.1975 0.08975 0.08827 -0.0464 0.9934 0.0062
-9.000 -0.1890 0.08554 0.08407 -0.0488 0.9909 0.0069
-8.750 -0.1799 0.08140 0.07993 -0.0517 0.9882 0.0073
-8.500 -0.1701 0.07725 0.07578 -0.0546 0.9856 0.0074
-8.250 -0.1593 0.07297 0.07150 -0.0578 0.9832 0.0076
-8.000 -0.1511 0.06878 0.06732 -0.0603 0.9779 0.0076
-7.750 -0.1404 0.06424 0.06278 -0.0638 0.9737 0.0077
-7.500 -0.1266 0.05924 0.05777 -0.0685 0.9704 0.0077
-7.250 -0.1139 0.05429 0.05282 -0.0733 0.9623 0.0077
-7.000 -0.1112 0.04719 0.04570 -0.0796 0.9483 0.0080
-6.750 -0.1012 0.04314 0.04161 -0.0837 0.9342 0.0084
-6.500 -0.0912 0.03980 0.03820 -0.0867 0.9204 0.0086
-6.250 -0.0808 0.03647 0.03481 -0.0896 0.9078 0.0088
-6.000 -0.0682 0.03306 0.03134 -0.0928 0.8969 0.0091
-5.750 -0.0533 0.02966 0.02787 -0.0962 0.8869 0.0098
-5.500 -0.0359 0.02561 0.02373 -0.1002 0.8776 0.0105
-5.250 -0.0153 0.02106 0.01907 -0.1047 0.8696 0.0110
-5.000 0.0084 0.01620 0.01404 -0.1092 0.8628 0.0117
-4.750 0.0354 0.01202 0.00962 -0.1124 0.8572 0.0126
-4.500 0.0603 0.00954 0.00682 -0.1141 0.8505 0.0130
-4.250 0.0842 0.00738 0.00432 -0.1151 0.8441 0.0130
-4.000 0.1084 0.00573 0.00232 -0.1157 0.8373 0.0131
-3.750 0.1396 0.01088 0.00618 -0.1214 0.8469 0.0084
-3.500 0.1653 0.00972 0.00485 -0.1209 0.8392 0.0089
-3.250 0.1918 0.00930 0.00433 -0.1207 0.8320 0.0102
-3.000 0.2183 0.00877 0.00369 -0.1203 0.8241 0.0111
-2.750 0.2449 0.00833 0.00316 -0.1199 0.8166 0.0118
-2.500 0.2718 0.00811 0.00286 -0.1196 0.8082 0.0128
-2.000 0.3247 0.00710 0.00155 -0.1187 0.7904 0.0163
-1.750 0.3515 0.00692 0.00129 -0.1184 0.7805 0.0193
-1.500 0.3781 0.00664 0.00105 -0.1180 0.7700 0.0526
-1.250 0.4041 0.00639 0.00100 -0.1178 0.7585 0.1363
-1.000 0.4307 0.00638 0.00096 -0.1175 0.7447 0.1514
-0.750 0.4568 0.00639 0.00091 -0.1171 0.7271 0.1655
-0.500 0.4827 0.00642 0.00088 -0.1166 0.7067 0.1784
-0.250 0.5084 0.00645 0.00087 -0.1162 0.6869 0.1951
0.000 0.5342 0.00647 0.00087 -0.1158 0.6699 0.2148
0.250 0.5602 0.00647 0.00089 -0.1155 0.6545 0.2513
0.500 0.5862 0.00646 0.00093 -0.1151 0.6402 0.2966
0.750 0.6121 0.00644 0.00097 -0.1148 0.6270 0.3499
1.000 0.6362 0.00610 0.00107 -0.1143 0.6149 0.5501
1.250 0.6745 0.00528 0.00119 -0.1169 0.6019 1.0000
1.500 0.7007 0.00538 0.00123 -0.1165 0.5913 1.0000
1.750 0.7270 0.00547 0.00128 -0.1162 0.5831 1.0000
2.250 0.7794 0.00567 0.00140 -0.1156 0.5659 1.0000
2.500 0.8055 0.00577 0.00149 -0.1152 0.5575 1.0000
2.750 0.8313 0.00590 0.00157 -0.1148 0.5454 1.0000
3.000 0.8560 0.00608 0.00166 -0.1142 0.5213 1.0000
3.250 0.8809 0.00627 0.00176 -0.1137 0.4972 1.0000
3.500 0.9053 0.00649 0.00188 -0.1131 0.4697 1.0000
3.750 0.9259 0.00701 0.00211 -0.1118 0.4013 1.0000
4.000 0.9375 0.00837 0.00272 -0.1091 0.2528 1.0000
4.250 0.9524 0.00949 0.00326 -0.1071 0.1478 1.0000
4.500 0.9642 0.01093 0.00412 -0.1044 0.0180 1.0000
4.750 0.9877 0.01129 0.00454 -0.1035 0.0138 1.0000
5.000 1.0106 0.01170 0.00504 -0.1025 0.0116 1.0000
5.250 1.0336 0.01208 0.00547 -0.1017 0.0106 1.0000
5.500 1.0558 0.01253 0.00598 -0.1006 0.0098 1.0000
5.750 1.0772 0.01305 0.00656 -0.0995 0.0087 1.0000
6.000 1.0968 0.01372 0.00729 -0.0980 0.0079 1.0000
6.250 1.1131 0.01466 0.00833 -0.0959 0.0074 1.0000
6.500 1.1195 0.01642 0.01022 -0.0921 0.0068 1.0000
6.750 1.1350 0.01739 0.01126 -0.0900 0.0065 1.0000
7.000 1.1526 0.01820 0.01212 -0.0882 0.0063 1.0000
7.250 1.1707 0.01896 0.01294 -0.0866 0.0060 1.0000
7.500 1.1883 0.01983 0.01387 -0.0849 0.0056 1.0000
7.750 1.2054 0.02099 0.01509 -0.0831 0.0052 1.0000
8.000 1.2243 0.02257 0.01674 -0.0816 0.0049 1.0000
8.750 1.2636 0.02921 0.02437 -0.0765 0.0066 1.0000
9.000 1.2721 0.03236 0.02775 -0.0740 0.0066 1.0000
9.250 1.2775 0.03550 0.03112 -0.0712 0.0066 1.0000
9.500 1.2792 0.03859 0.03443 -0.0680 0.0065 1.0000
9.750 1.2780 0.04125 0.03730 -0.0645 0.0065 1.0000
10.000 1.2735 0.04338 0.03962 -0.0607 0.0064 1.0000
10.250 1.2781 0.04149 0.03781 -0.0566 0.0059 1.0000
10.500 1.2707 0.04316 0.03963 -0.0525 0.0056 1.0000
10.750 1.2569 0.04571 0.04234 -0.0485 0.0054 1.0000
11.000 1.2412 0.04876 0.04556 -0.0452 0.0053 1.0000
11.250 1.2226 0.05256 0.04953 -0.0425 0.0052 1.0000
11.500 1.2033 0.05680 0.05394 -0.0406 0.0051 1.0000
11.750 1.1820 0.06176 0.05906 -0.0394 0.0051 1.0000
12.000 1.1593 0.06727 0.06473 -0.0391 0.0051 1.0000
12.250 1.1344 0.07348 0.07110 -0.0396 0.0052 1.0000
12.500 1.1087 0.08002 0.07778 -0.0410 0.0052 1.0000
12.750 1.0815 0.08691 0.08482 -0.0434 0.0053 1.0000
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Polar data table (+)
Polar graphs
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