XFOIL Version 6.94 Calculated polar for: GOE 406c 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 -0.2554 0.07544 0.06823 0.0380 1.0001 0.6943 -2.750 -0.2557 0.07380 0.06665 0.0297 1.0001 0.6441 -2.500 -0.2188 0.07090 0.06374 0.0149 1.0001 0.5826 -2.250 -0.1582 0.06810 0.06076 -0.0046 1.0001 0.5194 -2.000 -0.0967 0.06564 0.05813 -0.0209 1.0001 0.4730 -1.750 -0.0378 0.06383 0.05611 -0.0347 1.0001 0.4389 -1.500 0.0078 0.06257 0.05477 -0.0435 1.0001 0.4166 -1.250 0.0407 0.06220 0.05449 -0.0496 1.0001 0.4013 -0.750 0.4555 0.05321 0.04323 -0.1201 0.7422 0.3664 -0.500 0.5036 0.05280 0.04217 -0.1223 0.7042 0.3697 -0.250 0.5414 0.05285 0.04175 -0.1231 0.6756 0.3782 0.000 0.5732 0.05332 0.04195 -0.1235 0.6531 0.3920 0.250 0.6059 0.05370 0.04220 -0.1239 0.6367 0.4117 0.500 0.6299 0.05514 0.04379 -0.1247 0.6241 0.4372 0.750 0.6556 0.05611 0.04518 -0.1252 0.6146 0.4842 1.000 0.6661 0.05715 0.04754 -0.1239 0.6069 0.6039 1.250 0.6968 0.05849 0.04817 -0.1233 0.5985 0.9999 1.500 0.7012 0.06305 0.05243 -0.1238 0.5926 0.9999 1.750 0.7108 0.06703 0.05592 -0.1239 0.5869 0.9999 2.000 0.7384 0.06955 0.05757 -0.1238 0.5802 0.9999 2.250 0.7441 0.07399 0.06166 -0.1242 0.5780 0.9999 2.750 0.6871 0.08773 0.07579 -0.1244 0.5903 0.9999 3.000 0.6859 0.09269 0.08047 -0.1249 0.5956 0.9999 3.250 0.6582 0.09887 0.08678 -0.1249 0.6080 0.9999 3.500 0.6647 0.10352 0.09108 -0.1261 0.6163 0.9999 3.750 0.6450 0.10894 0.09652 -0.1265 0.6338 0.9999 4.000 0.6338 0.11396 0.10146 -0.1274 0.6541 0.9999 4.250 0.6080 0.11910 0.10673 -0.1282 0.6931 0.9999 5.000 0.4209 0.12548 0.11488 -0.1104 1.0001 0.9999