XFOIL Version 6.94 Calculated polar for: EPPLER 862 STRUT AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.0004 0.02130 0.01302 -0.0001 0.4465 0.4471 0.500 0.0577 0.02127 0.01301 -0.0008 0.4423 0.4516 1.000 0.1148 0.02114 0.01287 -0.0016 0.4376 0.4564 1.500 0.1738 0.02100 0.01280 -0.0025 0.4329 0.4616 2.000 0.2338 0.02102 0.01279 -0.0035 0.4284 0.4665 2.500 0.2909 0.02161 0.01336 -0.0043 0.4235 0.4717 3.000 0.3466 0.02175 0.01358 -0.0048 0.4195 0.4769 3.500 0.4022 0.02172 0.01358 -0.0053 0.4145 0.4819 4.000 0.4595 0.02169 0.01364 -0.0060 0.4101 0.4876 4.500 0.5192 0.02178 0.01369 -0.0070 0.4060 0.4935 5.000 0.5749 0.02242 0.01433 -0.0076 0.4012 0.4998 5.500 0.6249 0.02276 0.01478 -0.0073 0.3968 0.5048 6.000 0.6768 0.02279 0.01492 -0.0073 0.3918 0.5115 6.500 0.7328 0.02287 0.01506 -0.0077 0.3875 0.5188 7.000 0.7929 0.02310 0.01521 -0.0088 0.3832 0.5264 7.500 0.8356 0.02385 0.01609 -0.0077 0.3786 0.5327 8.000 0.8758 0.02421 0.01662 -0.0061 0.3733 0.5408 8.500 0.9244 0.02441 0.01691 -0.0056 0.3687 0.5501 9.000 0.9827 0.02454 0.01696 -0.0063 0.3644 0.5599 9.500 1.0102 0.02552 0.01810 -0.0033 0.3596 0.5691 10.000 1.0168 0.02667 0.01951 0.0023 0.3542 0.5794 10.500 1.0573 0.02727 0.02013 0.0032 0.3493 0.5916 11.000 1.1191 0.02736 0.02020 0.0020 0.3447 0.6079 11.500 1.0919 0.03124 0.02446 0.0071 0.3382 0.6204 12.000 1.1135 0.03326 0.02661 0.0079 0.3324 0.6374 12.500 1.1751 0.03301 0.02630 0.0071 0.3277 0.6602 13.000 1.1038 0.04203 0.03583 0.0097 0.3181 0.6763 13.500 1.1527 0.04234 0.03623 0.0094 0.3131 0.7086 14.000 1.0551 0.05585 0.05011 0.0097 0.3000 0.7255 14.500 1.1268 0.05367 0.04812 0.0098 0.2967 0.7846