XFOIL Version 6.94 Calculated polar for: pp13 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2416 0.06925 0.06173 -0.0094 0.9998 0.3759 -2.750 -0.2327 0.06663 0.05918 -0.0073 0.9998 0.4028 -2.500 -0.2242 0.06415 0.05678 -0.0052 0.9998 0.4414 -2.250 -0.2200 0.06140 0.05415 -0.0011 0.9998 0.4859 -1.750 -0.2263 0.05519 0.04822 0.0121 0.9998 0.6106 -1.500 -0.2283 0.05175 0.04491 0.0186 0.9998 0.6599 -1.250 -0.2264 0.04874 0.04199 0.0236 0.9998 0.7055 -1.000 -0.2217 0.04581 0.03915 0.0278 0.9998 0.7459 -0.750 -0.2116 0.04327 0.03666 0.0304 0.9998 0.7822 -0.500 -0.1967 0.04079 0.03421 0.0320 0.9998 0.8139 -0.250 -0.1700 0.03902 0.03242 0.0304 0.9998 0.8409 0.000 -0.1302 0.03764 0.03102 0.0255 0.9998 0.8618 0.250 -0.0736 0.03699 0.03034 0.0163 0.9998 0.8707 0.500 -0.0067 0.03804 0.03127 0.0036 0.9998 0.8507 0.750 0.0713 0.04174 0.03472 -0.0142 0.9998 0.7533 1.000 0.1444 0.04505 0.03774 -0.0309 0.9998 0.6174 1.250 0.2062 0.04724 0.03953 -0.0431 0.9998 0.4855 1.500 0.2578 0.04837 0.04026 -0.0509 0.9998 0.3881 1.750 0.3027 0.04906 0.04058 -0.0561 0.9998 0.3273 2.000 0.3458 0.04918 0.04063 -0.0606 0.9998 0.2988 2.500 0.8034 0.03767 0.02454 -0.1052 0.1978 0.2765 2.750 0.8492 0.03889 0.02547 -0.1082 0.1895 0.2890 3.000 0.9220 0.03951 0.02640 -0.1156 0.1824 0.3388 3.250 1.0162 0.04035 0.02755 -0.1259 0.1734 1.0002 3.500 1.0670 0.04200 0.02869 -0.1278 0.1704 1.0002 3.750 1.1089 0.04371 0.03035 -0.1287 0.1707 1.0002 4.000 1.1448 0.04552 0.03234 -0.1286 0.1720 1.0002 4.250 1.1766 0.04748 0.03462 -0.1281 0.1746 1.0002 4.500 1.2056 0.04988 0.03743 -0.1273 0.1784 1.0002 4.750 1.2329 0.05277 0.04068 -0.1264 0.1828 1.0002 5.000 1.2581 0.05603 0.04424 -0.1255 0.1861 1.0002