XFOIL Version 6.94 Calculated polar for: GOE 406c 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1033 0.07130 0.06609 0.0297 0.9999 0.9063 -2.750 -0.1442 0.07096 0.06593 0.0373 0.9999 0.8779 -2.500 -0.1898 0.07054 0.06568 0.0455 0.9999 0.8545 -2.250 -0.2366 0.06982 0.06513 0.0537 0.9999 0.8344 -2.000 -0.2777 0.06869 0.06416 0.0593 0.9999 0.8090 -1.750 -0.2999 0.06712 0.06273 0.0577 0.9999 0.7652 -1.500 0.4270 0.04472 0.03650 -0.1078 0.6977 0.2845 -1.250 0.4833 0.04269 0.03352 -0.1098 0.6309 0.2769 -1.000 0.5250 0.04170 0.03171 -0.1105 0.5897 0.2757 -0.750 0.5627 0.04108 0.03029 -0.1107 0.5611 0.2775 -0.500 0.5962 0.04086 0.02944 -0.1106 0.5385 0.2800 -0.250 0.6289 0.04064 0.02868 -0.1102 0.5209 0.2834 0.000 0.6584 0.04065 0.02848 -0.1098 0.5068 0.2911 0.250 0.6887 0.04096 0.02857 -0.1097 0.4969 0.3052 0.500 0.7174 0.04125 0.02892 -0.1096 0.4884 0.3270 0.750 0.7501 0.04119 0.02905 -0.1099 0.4817 0.3705 1.000 0.7630 0.04033 0.03022 -0.1072 0.4765 1.0001 1.250 0.7934 0.04212 0.03118 -0.1067 0.4701 1.0001 1.500 0.8261 0.04349 0.03157 -0.1063 0.4637 1.0001 1.750 0.8460 0.04575 0.03355 -0.1059 0.4585 1.0001 2.000 0.8614 0.04838 0.03609 -0.1055 0.4538 1.0001 2.250 0.8763 0.05118 0.03880 -0.1052 0.4505 1.0001 2.500 0.8839 0.05480 0.04245 -0.1048 0.4495 1.0001 2.750 0.8757 0.05995 0.04778 -0.1040 0.4505 1.0001 3.000 0.8426 0.06760 0.05576 -0.1033 0.4548 1.0001 3.250 0.8203 0.07462 0.06290 -0.1034 0.4594 1.0001 3.750 0.6879 0.09812 0.08740 -0.1094 0.4945 1.0001 4.000 0.6910 0.10326 0.09235 -0.1111 0.5046 1.0001 4.250 0.6825 0.10965 0.09870 -0.1136 0.5233 1.0001