XFOIL Version 6.94 Calculated polar for: manu4/30 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2188 0.07147 0.06539 0.0270 0.9310 0.8090 -2.750 -0.2742 0.06956 0.06361 0.0363 0.9328 0.7844 -2.500 -0.2974 0.06692 0.06100 0.0366 0.9150 0.7387 -2.250 -0.2507 0.06428 0.05819 0.0214 0.8493 0.6523 -2.000 -0.2113 0.06310 0.05684 0.0116 0.7657 0.5833 -1.750 -0.0453 0.05821 0.04925 -0.0205 0.5509 0.4794 -1.250 0.0624 0.05555 0.04495 -0.0353 0.5164 0.4287 -1.000 0.1256 0.05459 0.04307 -0.0439 0.5054 0.4108 -0.750 0.1835 0.05333 0.04128 -0.0508 0.4964 0.4040 -0.500 0.2430 0.05252 0.03987 -0.0578 0.4875 0.4076 -0.250 0.3137 0.05209 0.03864 -0.0670 0.4780 0.4173 0.000 0.3716 0.05138 0.03766 -0.0735 0.4710 0.4295 0.250 0.4398 0.05088 0.03679 -0.0820 0.4641 0.4553 0.500 0.5182 0.04994 0.03579 -0.0924 0.4600 0.5232 0.750 0.6937 0.04813 0.03501 -0.1240 0.4558 1.0001 1.000 0.7423 0.04971 0.03623 -0.1290 0.4559 1.0001 1.250 0.7818 0.05125 0.03757 -0.1320 0.4567 1.0001 1.500 0.8141 0.05280 0.03900 -0.1337 0.4582 1.0001 1.750 0.8408 0.05443 0.04058 -0.1343 0.4603 1.0001 2.000 0.8631 0.05621 0.04234 -0.1343 0.4628 1.0001 2.250 0.8822 0.05818 0.04431 -0.1337 0.4659 1.0001 2.500 0.8997 0.06034 0.04648 -0.1331 0.4692 1.0001 2.750 0.9151 0.06272 0.04885 -0.1321 0.4727 1.0001 3.000 0.9323 0.06529 0.05137 -0.1316 0.4762 1.0001 3.250 0.9289 0.06805 0.05429 -0.1282 0.4810 1.0001 3.500 0.8828 0.07251 0.05915 -0.1198 0.4899 1.0001 3.750 0.8602 0.07726 0.06401 -0.1155 0.4981 1.0001 4.000 0.8221 0.08292 0.06983 -0.1103 0.5082 1.0001 4.500 0.6275 0.10627 0.09383 -0.1030 0.5752 1.0001