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 = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4333 0.07208 0.06008 0.0424 0.9999 0.6552 -2.750 -0.4158 0.06990 0.05769 0.0385 0.9999 0.6390 -2.500 -0.3914 0.06736 0.05505 0.0352 0.9999 0.6291 -2.250 -0.3657 0.06541 0.05288 0.0311 0.9999 0.6211 -2.000 -0.3361 0.06330 0.05057 0.0276 0.9999 0.6172 -1.750 -0.3034 0.06133 0.04843 0.0237 0.9999 0.6141 -1.500 -0.2674 0.05962 0.04647 0.0194 0.9999 0.6110 -1.250 -0.2288 0.05809 0.04473 0.0150 0.9999 0.6094 -1.000 -0.1888 0.05664 0.04311 0.0106 0.9999 0.6130 -0.750 -0.1480 0.05514 0.04153 0.0067 0.9999 0.6252 -0.500 -0.1045 0.05355 0.03996 0.0025 0.9999 0.6448 -0.250 -0.0560 0.05192 0.03852 -0.0027 0.9999 0.6712 0.000 -0.0022 0.04990 0.03717 -0.0090 0.9999 0.7168 0.250 0.0750 0.04784 0.03707 -0.0224 0.9999 1.0001 0.500 0.0779 0.05138 0.04081 -0.0260 0.9999 1.0001 0.750 0.2581 0.05868 0.04563 -0.0695 0.8803 1.0001 1.000 0.3356 0.06044 0.04607 -0.0810 0.8254 1.0001 1.250 0.3872 0.06198 0.04679 -0.0867 0.7906 1.0001 1.500 0.4258 0.06379 0.04799 -0.0903 0.7678 1.0001 1.750 0.4461 0.06630 0.05008 -0.0917 0.7531 1.0001 2.000 0.4636 0.06904 0.05243 -0.0928 0.7436 1.0001 2.250 0.4763 0.07199 0.05508 -0.0933 0.7368 1.0001 2.500 0.4984 0.07476 0.05752 -0.0950 0.7309 1.0001 2.750 0.5071 0.07800 0.06048 -0.0953 0.7286 1.0001 3.000 0.5117 0.08141 0.06366 -0.0952 0.7280 1.0001 3.250 0.5178 0.08481 0.06685 -0.0953 0.7285 1.0001 3.500 0.5227 0.08827 0.07012 -0.0953 0.7304 1.0001 3.750 0.5138 0.09187 0.07357 -0.0938 0.7355 1.0001 4.000 0.5134 0.09534 0.07686 -0.0933 0.7410 1.0001 4.250 0.5211 0.09882 0.08015 -0.0938 0.7466 1.0001 4.500 0.5127 0.10211 0.08334 -0.0924 0.7561 1.0001 4.750 0.5156 0.10553 0.08659 -0.0924 0.7657 1.0001 5.000 0.5066 0.10860 0.08956 -0.0910 0.7792 1.0001