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 = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -1.250 0.4602 0.03307 0.02873 -0.0991 0.5722 0.0114 -1.000 0.4939 0.03143 0.02700 -0.1014 0.5673 0.0126 -0.750 0.5283 0.02984 0.02532 -0.1036 0.5619 0.0133 -0.500 0.5633 0.02858 0.02392 -0.1055 0.5572 0.0141 -0.250 0.6006 0.02781 0.02304 -0.1073 0.5523 0.0146 0.000 0.6348 0.02668 0.02182 -0.1088 0.5475 0.0148 0.250 0.6675 0.02542 0.02040 -0.1100 0.5430 0.0149 0.500 0.6991 0.02288 0.01778 -0.1118 0.5386 0.0156 0.750 0.7278 0.02155 0.01640 -0.1128 0.5337 0.0167 1.000 0.7584 0.02055 0.01527 -0.1136 0.5295 0.0181 1.250 0.7902 0.01960 0.01421 -0.1142 0.5251 0.0205 1.500 0.8241 0.01942 0.01391 -0.1142 0.5203 0.0228 2.500 0.9441 0.01519 0.00920 -0.1157 0.5028 0.0355 2.750 0.9720 0.01462 0.00852 -0.1158 0.4986 0.0374 3.000 1.0009 0.01397 0.00781 -0.1158 0.4940 0.0393 3.250 1.0334 0.01179 0.00522 -0.1153 0.4900 0.0224 3.500 1.0610 0.01086 0.00400 -0.1150 0.4859 0.0215 3.750 1.0880 0.01049 0.00356 -0.1147 0.4812 0.0229 4.250 1.1414 0.01040 0.00343 -0.1142 0.4706 0.0653 4.500 1.1682 0.01049 0.00359 -0.1140 0.4645 0.0812 4.750 1.1948 0.01052 0.00361 -0.1138 0.4591 0.0898 5.000 1.2214 0.01057 0.00373 -0.1136 0.4536 0.1005 5.250 1.2479 0.01062 0.00381 -0.1134 0.4472 0.1180 5.500 1.2743 0.01072 0.00395 -0.1132 0.4408 0.1401 5.750 1.3007 0.01079 0.00409 -0.1131 0.4339 0.1626 6.000 1.3268 0.01092 0.00424 -0.1128 0.4275 0.1848 6.250 1.3532 0.01097 0.00444 -0.1127 0.4199 0.2152 6.500 1.3804 0.01028 0.00465 -0.1129 0.4123 1.0000 6.750 1.4063 0.01044 0.00481 -0.1126 0.4034 1.0000 7.000 1.4321 0.01061 0.00501 -0.1124 0.3942 1.0000 7.250 1.4575 0.01081 0.00522 -0.1121 0.3843 1.0000 7.500 1.4826 0.01103 0.00544 -0.1118 0.3668 1.0000 7.750 1.5059 0.01149 0.00578 -0.1114 0.3293 1.0000 8.000 1.5283 0.01213 0.00624 -0.1109 0.2883 1.0000 8.250 1.5484 0.01314 0.00696 -0.1102 0.2351 1.0000 8.500 1.5649 0.01463 0.00804 -0.1093 0.1697 1.0000 8.750 1.5805 0.01617 0.00921 -0.1083 0.1141 1.0000 9.000 1.5937 0.01789 0.01059 -0.1069 0.0623 1.0000 9.250 1.6013 0.02012 0.01245 -0.1049 0.0141 1.0000 9.500 1.6159 0.02137 0.01372 -0.1034 0.0078 1.0000 9.750 1.6318 0.02235 0.01485 -0.1020 0.0070 1.0000 10.000 1.6452 0.02348 0.01612 -0.1004 0.0064 1.0000 10.250 1.6553 0.02478 0.01756 -0.0984 0.0060 1.0000 10.500 1.6592 0.02629 0.01920 -0.0957 0.0056 1.0000 10.750 1.6564 0.02800 0.02106 -0.0924 0.0054 1.0000 11.000 1.6499 0.03037 0.02361 -0.0897 0.0052 1.0000 11.250 1.6451 0.03304 0.02643 -0.0881 0.0051 1.0000 11.500 1.6380 0.03631 0.02986 -0.0869 0.0050 1.0000 11.750 1.6302 0.03997 0.03367 -0.0862 0.0049 1.0000 12.000 1.6228 0.04378 0.03763 -0.0858 0.0049 1.0000 12.250 1.6174 0.04744 0.04142 -0.0855 0.0049 1.0000 12.500 1.6094 0.05152 0.04565 -0.0853 0.0049 1.0000 12.750 1.6039 0.05534 0.04960 -0.0852 0.0049 1.0000 13.000 1.5984 0.05920 0.05358 -0.0851 0.0049 1.0000 13.250 1.5932 0.06308 0.05760 -0.0849 0.0049 1.0000 13.500 1.5894 0.06684 0.06148 -0.0849 0.0050 1.0000 13.750 1.5860 0.07065 0.06543 -0.0848 0.0050 1.0000 14.000 1.5838 0.07441 0.06934 -0.0849 0.0051 1.0000 14.250 1.5810 0.07835 0.07346 -0.0849 0.0053 1.0000 14.500 1.5776 0.08257 0.07790 -0.0846 0.0055 1.0000 14.750 1.5713 0.08747 0.08307 -0.0843 0.0059 1.0000 15.000 1.5613 0.09332 0.08920 -0.0847 0.0063 1.0000 15.250 1.5487 0.09977 0.09591 -0.0862 0.0067 1.0000 15.500 1.5344 0.10667 0.10303 -0.0885 0.0069 1.0000 15.750 1.5192 0.11399 0.11057 -0.0915 0.0071 1.0000