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 = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1766 0.05377 0.04597 -0.0145 0.3766 0.5242 -2.750 -0.1739 0.05199 0.04416 -0.0099 0.3697 0.5600 -2.500 -0.1670 0.04945 0.04158 -0.0055 0.3640 0.5873 -2.250 -0.1558 0.04739 0.03939 -0.0029 0.3588 0.6116 -2.000 -0.1364 0.04543 0.03728 -0.0027 0.3538 0.6249 -1.500 0.0751 0.04007 0.02809 -0.0487 0.3371 0.2121 -1.250 0.1204 0.03869 0.02615 -0.0519 0.3318 0.2153 -1.000 0.1698 0.03749 0.02472 -0.0561 0.3272 0.2198 -0.750 0.2291 0.03642 0.02326 -0.0623 0.3237 0.2251 -0.500 0.2967 0.03553 0.02192 -0.0702 0.3207 0.2356 -0.250 0.3777 0.03525 0.02134 -0.0814 0.3181 0.2604 0.000 0.4707 0.03479 0.02074 -0.0953 0.3161 0.3443 0.250 0.5333 0.03463 0.02095 -0.1029 0.3154 0.4495 0.500 0.5753 0.03494 0.02154 -0.1060 0.3153 0.5031 0.750 0.6127 0.03523 0.02229 -0.1081 0.3155 0.5699 1.000 0.7224 0.03622 0.02408 -0.1265 0.3164 1.0001 1.250 0.7531 0.03720 0.02492 -0.1270 0.3171 1.0001 1.500 0.7810 0.03820 0.02585 -0.1269 0.3176 1.0001 1.750 0.8073 0.03932 0.02693 -0.1266 0.3177 1.0001 2.000 0.8323 0.04045 0.02803 -0.1260 0.3180 1.0001 2.250 0.8558 0.04156 0.02922 -0.1251 0.3192 1.0001 2.500 0.8782 0.04282 0.03059 -0.1240 0.3210 1.0001 2.750 0.9000 0.04425 0.03220 -0.1230 0.3248 1.0001 3.000 0.9210 0.04587 0.03392 -0.1219 0.3284 1.0001 3.250 0.9421 0.04764 0.03574 -0.1209 0.3318 1.0001 3.500 0.9639 0.04970 0.03778 -0.1202 0.3347 1.0001 3.750 0.9782 0.05090 0.03944 -0.1180 0.3434 1.0001 4.000 0.9930 0.05318 0.04195 -0.1164 0.3512 1.0001 4.250 1.0144 0.05582 0.04456 -0.1159 0.3561 1.0001 4.500 1.0181 0.05795 0.04726 -0.1132 0.3738 1.0001 5.000 1.0343 0.06496 0.05493 -0.1116 0.4224 1.0001