XFOIL Version 6.94 Calculated polar for: manu4/36 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3265 0.06287 0.05665 0.0492 0.9999 0.8517 -2.750 -0.3903 0.06043 0.05435 0.0614 0.9999 0.8307 -2.500 -0.4515 0.05758 0.05160 0.0715 0.9999 0.8062 -2.250 -0.4671 0.05550 0.04950 0.0685 0.9999 0.7497 -2.000 -0.3532 0.05775 0.05122 0.0307 0.9999 0.5472 -1.750 -0.0967 0.04921 0.04118 -0.0226 0.7617 0.3377 -1.500 0.0204 0.04526 0.03436 -0.0381 0.4880 0.3203 -1.250 0.0554 0.04417 0.03260 -0.0398 0.4706 0.3156 -1.000 0.0979 0.04306 0.03087 -0.0428 0.4575 0.3143 -0.750 0.1569 0.04227 0.02918 -0.0490 0.4461 0.3204 -0.500 0.2157 0.04136 0.02780 -0.0551 0.4375 0.3371 -0.250 0.2874 0.04060 0.02641 -0.0639 0.4282 0.3614 0.000 0.3757 0.03943 0.02494 -0.0761 0.4186 0.4222 0.250 0.4594 0.03693 0.02346 -0.0871 0.4095 0.6410 0.500 0.6025 0.03806 0.02440 -0.1122 0.4021 1.0001 0.750 0.6498 0.03902 0.02502 -0.1162 0.4011 1.0001 1.000 0.6911 0.03998 0.02577 -0.1190 0.4006 1.0001 1.250 0.7287 0.04103 0.02666 -0.1210 0.4005 1.0001 1.500 0.7633 0.04215 0.02769 -0.1225 0.4009 1.0001 1.750 0.7953 0.04340 0.02886 -0.1235 0.4015 1.0001 2.000 0.8239 0.04457 0.03007 -0.1238 0.4026 1.0001 2.250 0.8479 0.04565 0.03128 -0.1234 0.4047 1.0001 2.500 0.8707 0.04693 0.03272 -0.1228 0.4073 1.0001 2.750 0.8926 0.04838 0.03433 -0.1222 0.4104 1.0001 3.000 0.9136 0.05001 0.03609 -0.1215 0.4137 1.0001 3.250 0.9330 0.05179 0.03797 -0.1206 0.4167 1.0001 3.500 0.9517 0.05371 0.03998 -0.1197 0.4192 1.0001 3.750 0.9700 0.05578 0.04212 -0.1187 0.4214 1.0001 4.000 0.9887 0.05803 0.04441 -0.1179 0.4234 1.0001 4.250 0.9906 0.05994 0.04670 -0.1149 0.4284 1.0001 4.500 0.9867 0.06270 0.04975 -0.1114 0.4341 1.0001 4.750 0.9891 0.06573 0.05292 -0.1091 0.4398 1.0001 5.000 1.0025 0.06882 0.05605 -0.1084 0.4446 1.0001