XFOIL Version 6.96 Calculated polar for: EPPLER 376 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 ------ -------- --------- --------- -------- -------- -------- 0.250 0.6690 0.02305 0.01877 -0.1105 0.5069 0.0098 0.500 0.7014 0.02170 0.01732 -0.1116 0.5026 0.0098 0.750 0.7334 0.02040 0.01591 -0.1126 0.4984 0.0098 3.250 1.0366 0.00901 0.00310 -0.1153 0.4567 0.0157 3.500 1.0640 0.00865 0.00269 -0.1151 0.4515 0.0146 3.750 1.0909 0.00835 0.00227 -0.1148 0.4457 0.0147 4.000 1.1180 0.00820 0.00204 -0.1146 0.4407 0.0196 4.250 1.1451 0.00820 0.00208 -0.1144 0.4353 0.0487 4.500 1.1717 0.00832 0.00219 -0.1143 0.4292 0.0597 4.750 1.1987 0.00836 0.00227 -0.1141 0.4230 0.0699 5.000 1.2254 0.00845 0.00238 -0.1140 0.4165 0.0816 5.250 1.2522 0.00852 0.00248 -0.1139 0.4104 0.0990 5.500 1.2788 0.00862 0.00261 -0.1137 0.4035 0.1175 5.750 1.3054 0.00872 0.00273 -0.1136 0.3966 0.1339 6.000 1.3317 0.00886 0.00289 -0.1134 0.3884 0.1522 6.250 1.3582 0.00897 0.00307 -0.1133 0.3799 0.1708 6.500 1.3843 0.00913 0.00325 -0.1131 0.3709 0.1928 7.250 1.4591 0.00934 0.00413 -0.1126 0.2910 1.0000 7.500 1.4807 0.01027 0.00471 -0.1121 0.2332 1.0000 7.750 1.5018 0.01125 0.00542 -0.1116 0.1824 1.0000 8.000 1.5223 0.01227 0.00617 -0.1110 0.1355 1.0000 8.250 1.5408 0.01356 0.00710 -0.1102 0.0827 1.0000 8.500 1.5559 0.01525 0.00838 -0.1090 0.0267 1.0000 8.750 1.5728 0.01657 0.00955 -0.1078 0.0051 1.0000 9.000 1.5943 0.01717 0.01024 -0.1070 0.0044 1.0000 9.250 1.6150 0.01781 0.01097 -0.1062 0.0041 1.0000 9.500 1.6346 0.01856 0.01181 -0.1053 0.0038 1.0000 9.750 1.6523 0.01946 0.01281 -0.1042 0.0034 1.0000 10.000 1.6675 0.02056 0.01403 -0.1027 0.0031 1.0000 10.250 1.6795 0.02186 0.01547 -0.1009 0.0029 1.0000 10.500 1.6868 0.02341 0.01718 -0.0985 0.0028 1.0000 10.750 1.6889 0.02498 0.01889 -0.0954 0.0027 1.0000 11.000 1.6827 0.02674 0.02078 -0.0916 0.0027 1.0000 11.250 1.6797 0.02871 0.02287 -0.0891 0.0027 1.0000 11.500 1.6749 0.03124 0.02553 -0.0872 0.0027 1.0000 11.750 1.6701 0.03416 0.02858 -0.0860 0.0027 1.0000 12.000 1.6657 0.03729 0.03184 -0.0852 0.0027 1.0000 12.250 1.6639 0.04031 0.03497 -0.0847 0.0028 1.0000 12.500 1.6585 0.04388 0.03867 -0.0843 0.0028 1.0000 12.750 1.6553 0.04729 0.04220 -0.0840 0.0029 1.0000 13.000 1.6507 0.05093 0.04597 -0.0838 0.0029 1.0000 13.250 1.6444 0.05487 0.05006 -0.0836 0.0030 1.0000 13.500 1.6377 0.05890 0.05425 -0.0834 0.0031 1.0000 13.750 1.6307 0.06308 0.05857 -0.0833 0.0032 1.0000 14.000 1.6251 0.06715 0.06278 -0.0834 0.0032 1.0000 14.250 1.6202 0.07134 0.06709 -0.0838 0.0032 1.0000 14.500 1.6145 0.07576 0.07165 -0.0843 0.0032 1.0000 14.750 1.6064 0.08072 0.07682 -0.0846 0.0033 1.0000 15.000 1.5947 0.08655 0.08290 -0.0850 0.0036 1.0000 15.250 1.5820 0.09275 0.08931 -0.0863 0.0038 1.0000 15.500 1.5665 0.09974 0.09651 -0.0883 0.0041 1.0000 16.500 1.3165 0.13096 0.12863 -0.0936 0.0052 1.0000 16.750 1.2602 0.14524 0.14343 -0.1009 0.0070 1.0000