XFOIL Version 6.94 Calculated polar for: nabac 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2053 0.05592 0.04033 -0.0042 0.9999 1.0001 -2.750 -0.2004 0.05355 0.03705 -0.0078 0.9999 1.0001 -2.500 -0.1693 0.05271 0.03404 -0.0113 0.9999 1.0001 -2.250 -0.1384 0.05248 0.03218 -0.0116 0.9999 1.0001 -2.000 -0.1107 0.05234 0.03092 -0.0109 0.9999 1.0001 -1.750 -0.0846 0.05224 0.03000 -0.0099 0.9999 1.0001 -1.500 -0.0593 0.05216 0.02931 -0.0089 0.9999 1.0001 -1.250 -0.0342 0.05212 0.02880 -0.0078 0.9999 1.0001 -1.000 -0.0092 0.05210 0.02845 -0.0068 0.9999 1.0001 -0.750 0.0159 0.05211 0.02825 -0.0058 0.9999 1.0001 -0.500 0.0414 0.05214 0.02820 -0.0050 0.9999 1.0001 -0.250 0.0675 0.05219 0.02831 -0.0043 0.9999 1.0001 0.000 0.0944 0.05229 0.02861 -0.0039 0.9999 1.0001 0.250 0.1220 0.05247 0.02905 -0.0041 0.9999 1.0001 0.500 0.1493 0.05283 0.02993 -0.0049 0.9999 1.0001 0.750 0.1723 0.05364 0.03139 -0.0065 0.9999 1.0001 1.000 0.1810 0.05541 0.03363 -0.0079 0.9999 1.0001 1.250 0.1751 0.05816 0.03627 -0.0083 0.9999 1.0001 1.500 0.1684 0.06115 0.03884 -0.0086 0.9999 1.0001 1.750 0.1657 0.06398 0.04121 -0.0092 0.9999 1.0001 2.000 0.1663 0.06664 0.04344 -0.0099 0.9999 1.0001 2.250 0.1691 0.06919 0.04557 -0.0107 0.9999 1.0001 2.500 0.1736 0.07168 0.04764 -0.0115 0.9999 1.0001 2.750 0.1792 0.07411 0.04968 -0.0123 0.9999 1.0001 3.000 0.1858 0.07652 0.05171 -0.0131 0.9999 1.0001 3.250 0.1931 0.07892 0.05372 -0.0139 0.9999 1.0001 3.500 0.2010 0.08130 0.05574 -0.0147 0.9999 1.0001 3.750 0.2095 0.08370 0.05778 -0.0155 0.9999 1.0001 4.000 0.2183 0.08610 0.05984 -0.0163 0.9999 1.0001 4.250 0.2275 0.08851 0.06192 -0.0171 0.9999 1.0001 4.500 0.2368 0.09093 0.06401 -0.0179 0.9999 1.0001 4.750 0.2465 0.09336 0.06613 -0.0186 0.9999 1.0001 5.000 0.2563 0.09580 0.06818 -0.0194 0.9999 1.0001