XFOIL Version 6.94 Calculated polar for: manu04 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4254 0.06895 0.05660 0.0377 0.9999 0.6418 -2.750 -0.4045 0.06675 0.05415 0.0335 0.9999 0.6268 -2.500 -0.3782 0.06428 0.05149 0.0300 0.9999 0.6177 -2.250 -0.3507 0.06214 0.04913 0.0263 0.9999 0.6126 -2.000 -0.3203 0.06012 0.04688 0.0227 0.9999 0.6115 -1.750 -0.2870 0.05823 0.04474 0.0190 0.9999 0.6119 -1.500 -0.2509 0.05654 0.04276 0.0150 0.9999 0.6128 -1.250 -0.2118 0.05501 0.04097 0.0107 0.9999 0.6148 -1.000 -0.1705 0.05360 0.03932 0.0063 0.9999 0.6217 -0.750 -0.1294 0.05199 0.03762 0.0026 0.9999 0.6401 -0.500 -0.0854 0.05017 0.03585 -0.0015 0.9999 0.6696 -0.250 -0.0339 0.04800 0.03409 -0.0069 0.9999 0.7142 0.000 0.0594 0.04497 0.03266 -0.0214 0.9999 1.0001 0.250 0.1369 0.04630 0.03308 -0.0348 0.9999 1.0001 0.500 0.1814 0.04819 0.03450 -0.0417 0.9999 1.0001 0.750 0.1723 0.05221 0.03873 -0.0430 0.9999 1.0001 1.000 0.2874 0.05786 0.04303 -0.0700 0.8996 1.0001 1.250 0.3556 0.05944 0.04375 -0.0797 0.8413 1.0001 1.500 0.3958 0.06112 0.04488 -0.0839 0.8068 1.0001 1.750 0.4370 0.06282 0.04608 -0.0879 0.7810 1.0001 2.000 0.4648 0.06503 0.04793 -0.0901 0.7645 1.0001 2.250 0.4752 0.06790 0.05051 -0.0903 0.7551 1.0001 2.500 0.4950 0.07061 0.05294 -0.0917 0.7468 1.0001 2.750 0.5117 0.07350 0.05556 -0.0928 0.7413 1.0001 3.000 0.5150 0.07680 0.05866 -0.0924 0.7389 1.0001 3.250 0.5202 0.08007 0.06173 -0.0924 0.7382 1.0001 3.500 0.5256 0.08339 0.06487 -0.0924 0.7385 1.0001 3.750 0.5323 0.08672 0.06803 -0.0926 0.7403 1.0001 4.000 0.5413 0.09012 0.07125 -0.0931 0.7427 1.0001 4.250 0.5291 0.09351 0.07454 -0.0912 0.7497 1.0001 4.500 0.5295 0.09689 0.07778 -0.0909 0.7563 1.0001 4.750 0.5409 0.10040 0.08112 -0.0919 0.7626 1.0001 5.000 0.5254 0.10337 0.08402 -0.0896 0.7753 1.0001