GOE 370 AIRFOIL (goe370-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 370 AIRFOIL (goe370-il) Reynolds number: 500,000 Max Cl/Cd: 128.11 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe370-il-500000.txt Download as CSV file: xf-goe370-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 370 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.3080 0.09768 0.09551 -0.0281 1.0000 0.0130
-7.750 -0.3108 0.09595 0.09383 -0.0273 1.0000 0.0131
-7.500 -0.3203 0.09490 0.09285 -0.0251 1.0000 0.0131
-7.250 -0.3102 0.09194 0.08991 -0.0283 0.9976 0.0131
-7.000 -0.2809 0.08719 0.08515 -0.0367 0.9930 0.0132
-6.750 -0.2532 0.08255 0.08050 -0.0442 0.9876 0.0133
-6.500 -0.2222 0.07768 0.07560 -0.0522 0.9834 0.0133
-6.250 -0.2110 0.07207 0.07001 -0.0534 0.9770 0.0142
-6.000 -0.1823 0.06830 0.06622 -0.0592 0.9716 0.0149
-5.750 -0.1524 0.06450 0.06238 -0.0655 0.9634 0.0163
-5.500 -0.0998 0.06021 0.05801 -0.0782 0.9568 0.0195
-5.250 -0.0574 0.05633 0.05402 -0.0865 0.9464 0.0199
-5.000 -0.0196 0.05182 0.04942 -0.0935 0.9354 0.0199
-4.750 0.0158 0.04745 0.04493 -0.0994 0.9227 0.0200
-4.500 0.0431 0.04040 0.03772 -0.1054 0.9072 0.0209
-4.250 0.0671 0.03792 0.03512 -0.1071 0.8936 0.0218
-4.000 0.0920 0.03591 0.03299 -0.1084 0.8812 0.0233
-3.750 0.1270 0.03402 0.03091 -0.1107 0.8704 0.0295
-3.000 0.2097 0.01972 0.01561 -0.1153 0.8436 0.0264
-2.750 0.2356 0.01430 0.00938 -0.1150 0.8361 0.0270
-2.500 0.2628 0.01398 0.00882 -0.1147 0.8291 0.0298
-2.250 0.2876 0.01147 0.00578 -0.1142 0.8221 0.0334
-2.000 0.3141 0.01082 0.00495 -0.1139 0.8156 0.0359
-1.750 0.3407 0.01040 0.00440 -0.1135 0.8088 0.0394
-1.500 0.3675 0.00997 0.00382 -0.1132 0.8029 0.0413
-1.250 0.3940 0.00955 0.00329 -0.1128 0.7964 0.0421
-1.000 0.4208 0.00925 0.00288 -0.1124 0.7907 0.0426
-0.750 0.4467 0.00877 0.00234 -0.1119 0.7839 0.0444
-0.500 0.4730 0.00848 0.00199 -0.1115 0.7778 0.0464
-0.250 0.4994 0.00830 0.00180 -0.1111 0.7712 0.0506
0.000 0.5260 0.00819 0.00165 -0.1107 0.7650 0.0573
0.250 0.5521 0.00802 0.00160 -0.1103 0.7576 0.0909
0.500 0.5784 0.00802 0.00163 -0.1099 0.7492 0.1295
0.750 0.6044 0.00800 0.00163 -0.1094 0.7387 0.1459
1.000 0.6303 0.00800 0.00162 -0.1090 0.7281 0.1623
1.250 0.6562 0.00797 0.00160 -0.1085 0.7180 0.1758
1.500 0.6822 0.00795 0.00160 -0.1081 0.7084 0.1908
1.750 0.7081 0.00787 0.00165 -0.1077 0.6987 0.2217
2.000 0.7316 0.00744 0.00180 -0.1071 0.6894 0.4774
2.250 0.7809 0.00652 0.00185 -0.1120 0.6790 1.0000
2.500 0.8062 0.00660 0.00189 -0.1114 0.6675 1.0000
2.750 0.8314 0.00668 0.00195 -0.1108 0.6556 1.0000
3.000 0.8566 0.00678 0.00205 -0.1102 0.6433 1.0000
3.250 0.8814 0.00688 0.00214 -0.1095 0.6284 1.0000
3.500 0.9039 0.00706 0.00219 -0.1083 0.5959 1.0000
3.750 0.9256 0.00735 0.00231 -0.1070 0.5594 1.0000
4.000 0.9474 0.00768 0.00249 -0.1058 0.5261 1.0000
4.250 0.9694 0.00803 0.00274 -0.1046 0.4904 1.0000
4.500 0.9851 0.00883 0.00306 -0.1024 0.3803 1.0000
4.750 0.9776 0.01210 0.00453 -0.0969 0.0538 1.0000
5.000 0.9968 0.01292 0.00527 -0.0952 0.0255 1.0000
5.250 1.0158 0.01377 0.00627 -0.0934 0.0210 1.0000
5.500 1.0375 0.01427 0.00682 -0.0923 0.0190 1.0000
5.750 1.0576 0.01492 0.00754 -0.0908 0.0171 1.0000
6.000 1.0761 0.01572 0.00843 -0.0891 0.0157 1.0000
6.250 1.0922 0.01673 0.00952 -0.0870 0.0146 1.0000
6.500 1.1034 0.01831 0.01119 -0.0841 0.0137 1.0000
6.750 1.1171 0.01999 0.01294 -0.0816 0.0135 1.0000
7.000 1.1364 0.02344 0.01645 -0.0802 0.0128 1.0000
7.250 1.1221 0.01194 0.00539 -0.0742 0.0120 1.0000
7.500 1.1454 0.01404 0.00761 -0.0733 0.0117 1.0000
7.750 1.1698 0.01695 0.01074 -0.0726 0.0116 1.0000
10.750 1.2370 0.07175 0.06818 -0.0448 0.0127 1.0000
11.000 1.2189 0.07463 0.07121 -0.0419 0.0127 1.0000
11.250 1.2004 0.07804 0.07477 -0.0402 0.0127 1.0000
11.500 1.1814 0.08172 0.07860 -0.0394 0.0127 1.0000
11.750 1.1630 0.08631 0.08332 -0.0397 0.0127 1.0000
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Polar data table (+)
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