XFOIL Version 6.94 Calculated polar for: GOA 3 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 -0.1480 0.02589 0.01384 -0.0322 0.9999 0.1315 -2.750 -0.1216 0.02427 0.01243 -0.0311 0.9999 0.1537 -2.500 -0.0941 0.02188 0.01094 -0.0307 0.9999 0.2587 -2.250 -0.0754 0.01867 0.01022 -0.0267 0.9999 1.0001 -2.000 -0.0501 0.01874 0.00988 -0.0254 0.9999 1.0001 -1.750 -0.0252 0.01883 0.00972 -0.0242 0.9999 1.0001 -1.500 0.0000 0.01893 0.00969 -0.0230 0.9999 1.0001 -1.250 0.0258 0.01902 0.00975 -0.0219 0.9999 1.0001 -0.750 0.1554 0.02217 0.00959 -0.0309 0.2668 1.0001 -0.500 0.1796 0.02313 0.01011 -0.0296 0.2498 1.0001 -0.250 0.2049 0.02411 0.01069 -0.0286 0.2324 1.0001 0.000 0.2321 0.02509 0.01134 -0.0279 0.2181 1.0001 0.250 0.2619 0.02616 0.01215 -0.0276 0.2118 1.0001 0.500 0.2944 0.02734 0.01318 -0.0275 0.2093 1.0001 0.750 0.3280 0.02846 0.01432 -0.0276 0.2084 1.0001 1.000 0.3616 0.02974 0.01567 -0.0278 0.2088 1.0001 1.250 0.3942 0.03118 0.01721 -0.0280 0.2101 1.0001 1.500 0.4264 0.03238 0.01880 -0.0279 0.2141 1.0001 1.750 0.4574 0.03404 0.02075 -0.0280 0.2194 1.0001 2.000 0.4869 0.03604 0.02290 -0.0281 0.2246 1.0001 2.250 0.5186 0.03781 0.02519 -0.0284 0.2341 1.0001 2.500 0.5474 0.04036 0.02787 -0.0287 0.2423 1.0001 2.750 0.5787 0.04271 0.03073 -0.0296 0.2578 1.0001 3.000 0.6126 0.04520 0.03398 -0.0317 0.2835 1.0001 3.500 0.6898 0.05256 0.04287 -0.0438 0.3900 1.0001 4.000 0.6514 0.06864 0.05985 -0.0868 0.6619 1.0001 4.250 0.5821 0.07320 0.06421 -0.0886 0.7671 1.0001 4.750 0.4062 0.06979 0.06041 -0.0620 0.9999 1.0001 5.000 0.4178 0.07279 0.06336 -0.0633 0.9999 1.0001