XFOIL Version 6.94 Calculated polar for: manu02 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0533 0.06168 0.05363 -0.0012 0.4917 0.8189 -2.750 -0.1297 0.05971 0.05211 0.0134 0.5044 0.7996 -2.500 -0.1722 0.05771 0.05030 0.0196 0.5079 0.7638 -2.250 -0.1834 0.05700 0.04929 0.0173 0.5007 0.7055 -2.000 -0.1185 0.05843 0.04981 -0.0016 0.4795 0.5916 -1.750 -0.0368 0.05900 0.04932 -0.0214 0.4617 0.4742 -1.500 0.0272 0.05750 0.04702 -0.0323 0.4519 0.4153 -1.250 0.0860 0.05594 0.04483 -0.0403 0.4454 0.3910 -1.000 0.1429 0.05474 0.04310 -0.0474 0.4396 0.3751 -0.750 0.1990 0.05322 0.04115 -0.0538 0.4351 0.3653 -0.500 0.2628 0.05235 0.03959 -0.0616 0.4309 0.3575 -0.250 0.3306 0.05144 0.03815 -0.0700 0.4275 0.3592 0.000 0.4022 0.05092 0.03715 -0.0791 0.4248 0.3745 0.250 0.4796 0.05053 0.03638 -0.0895 0.4230 0.4020 0.500 0.5733 0.04976 0.03553 -0.1033 0.4219 0.4660 0.750 0.6518 0.04811 0.03465 -0.1136 0.4219 0.6337 1.000 0.7763 0.04888 0.03618 -0.1352 0.4227 1.0001 1.250 0.8144 0.05041 0.03752 -0.1379 0.4232 1.0001 1.500 0.8462 0.05198 0.03895 -0.1392 0.4237 1.0001 1.750 0.8735 0.05360 0.04048 -0.1397 0.4241 1.0001 2.000 0.8971 0.05531 0.04212 -0.1395 0.4244 1.0001 2.250 0.9149 0.05701 0.04383 -0.1383 0.4258 1.0001 2.500 0.9284 0.05884 0.04573 -0.1365 0.4282 1.0001 2.750 0.9387 0.06092 0.04790 -0.1344 0.4319 1.0001 3.000 0.9472 0.06329 0.05036 -0.1321 0.4366 1.0001 3.250 0.9562 0.06596 0.05308 -0.1302 0.4415 1.0001 3.500 0.9720 0.06880 0.05586 -0.1294 0.4458 1.0001 3.750 0.9549 0.07171 0.05905 -0.1241 0.4532 1.0001 4.000 0.9007 0.07691 0.06462 -0.1153 0.4651 1.0001 4.250 0.9083 0.08087 0.06851 -0.1146 0.4729 1.0001 4.500 0.7838 0.09143 0.07954 -0.1031 0.4994 1.0001