XFOIL Version 6.94 Calculated polar for: Eif385mod 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1166 0.04407 0.02879 -0.0191 0.9993 0.5579 -2.750 -0.0894 0.04341 0.02816 -0.0197 0.9993 0.5685 -2.500 -0.0614 0.04294 0.02771 -0.0206 0.9993 0.5850 -2.250 -0.0356 0.04242 0.02741 -0.0207 0.9993 0.6107 -2.000 -0.0109 0.04185 0.02727 -0.0205 0.9993 0.6463 -1.750 0.0127 0.04111 0.02730 -0.0196 0.9993 0.7121 -1.500 0.0390 0.04054 0.02711 -0.0210 0.9993 1.0007 -1.250 0.0700 0.04141 0.02735 -0.0234 0.9993 1.0007 -1.000 0.0959 0.04234 0.02779 -0.0243 0.9993 1.0007 -0.750 0.1186 0.04331 0.02840 -0.0246 0.9993 1.0007 -0.500 0.1393 0.04435 0.02916 -0.0245 0.9993 1.0007 -0.250 0.1582 0.04549 0.03008 -0.0243 0.9993 1.0007 0.000 0.1755 0.04675 0.03119 -0.0241 0.9993 1.0007 0.250 0.1911 0.04820 0.03253 -0.0239 0.9993 1.0007 0.500 0.2044 0.04989 0.03417 -0.0238 0.9993 1.0007 0.750 0.2149 0.05197 0.03625 -0.0239 0.9993 1.0007 1.000 0.2213 0.05460 0.03892 -0.0244 0.9993 1.0007 1.250 0.2230 0.05787 0.04225 -0.0254 0.9993 1.0007 1.500 0.2233 0.06147 0.04589 -0.0267 0.9993 1.0007 1.750 0.2855 0.06587 0.05018 -0.0395 0.9651 1.0007 2.000 0.3513 0.06914 0.05333 -0.0507 0.9165 1.0007 2.250 0.3924 0.07213 0.05622 -0.0566 0.8884 1.0007 2.500 0.4168 0.07487 0.05887 -0.0593 0.8675 1.0007 2.750 0.4500 0.07765 0.06157 -0.0631 0.8477 1.0007 3.000 0.4664 0.08031 0.06417 -0.0642 0.8326 1.0007 3.250 0.4849 0.08310 0.06691 -0.0656 0.8206 1.0007 3.500 0.4903 0.08607 0.06984 -0.0655 0.8162 1.0007 3.750 0.4956 0.08919 0.07293 -0.0656 0.8160 1.0007 4.000 0.4998 0.09239 0.07609 -0.0656 0.8191 1.0007 4.250 0.4987 0.09555 0.07924 -0.0651 0.8269 1.0007 4.500 0.5046 0.09893 0.08259 -0.0656 0.8350 1.0007 4.750 0.4893 0.10170 0.08537 -0.0631 0.8554 1.0007 5.000 0.4675 0.10449 0.08814 -0.0598 0.8949 1.0007