XFOIL Version 6.94 Calculated polar for: manu4/33 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 -0.4193 0.07075 0.06359 0.0629 0.9999 0.7930 -2.750 -0.4503 0.06826 0.06119 0.0658 0.9999 0.7563 -2.500 -0.4538 0.06647 0.05937 0.0611 0.9999 0.7022 -2.250 -0.4098 0.06628 0.05900 0.0459 0.9999 0.6194 -2.000 -0.3520 0.06490 0.05754 0.0321 0.9999 0.5464 -1.750 -0.2997 0.06327 0.05584 0.0218 0.9999 0.4963 -1.500 -0.0984 0.05621 0.04819 -0.0167 0.8375 0.4255 -0.750 0.1465 0.04745 0.03533 -0.0457 0.5318 0.4068 -0.500 0.2144 0.04668 0.03368 -0.0541 0.5171 0.4154 -0.250 0.2854 0.04591 0.03223 -0.0630 0.5057 0.4300 0.000 0.3631 0.04519 0.03092 -0.0732 0.4958 0.4718 0.250 0.4389 0.04352 0.02945 -0.0826 0.4864 0.5678 0.500 0.6034 0.04279 0.02897 -0.1119 0.4709 1.0001 0.750 0.6505 0.04402 0.02977 -0.1162 0.4669 1.0001 1.000 0.6931 0.04525 0.03066 -0.1195 0.4651 1.0001 1.250 0.7321 0.04651 0.03169 -0.1221 0.4644 1.0001 1.500 0.7671 0.04782 0.03287 -0.1240 0.4644 1.0001 1.750 0.7987 0.04922 0.03419 -0.1253 0.4650 1.0001 2.000 0.8277 0.05073 0.03566 -0.1262 0.4661 1.0001 2.250 0.8544 0.05234 0.03727 -0.1268 0.4676 1.0001 2.500 0.8792 0.05410 0.03906 -0.1270 0.4694 1.0001 2.750 0.9022 0.05603 0.04100 -0.1271 0.4715 1.0001 3.000 0.9248 0.05813 0.04310 -0.1271 0.4737 1.0001 3.250 0.9264 0.06017 0.04543 -0.1241 0.4778 1.0001 3.500 0.9224 0.06291 0.04842 -0.1207 0.4831 1.0001 3.750 0.9169 0.06615 0.05182 -0.1175 0.4888 1.0001 4.000 0.9194 0.06954 0.05526 -0.1157 0.4942 1.0001 4.250 0.9009 0.07374 0.05963 -0.1119 0.5015 1.0001 4.500 0.8241 0.08149 0.06764 -0.1037 0.5166 1.0001