XFOIL Version 6.94 Calculated polar for: manu4/38 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -2.750 -0.0743 0.04698 0.03762 -0.0461 0.3943 0.2720 -2.500 -0.0505 0.04400 0.03397 -0.0465 0.3824 0.2346 -2.250 -0.0289 0.04198 0.03127 -0.0456 0.3722 0.2203 -2.000 -0.0062 0.04043 0.02940 -0.0446 0.3622 0.2163 -1.750 0.0212 0.03890 0.02735 -0.0444 0.3536 0.2127 -1.500 0.0542 0.03758 0.02556 -0.0452 0.3471 0.2124 -1.250 0.0923 0.03655 0.02416 -0.0469 0.3420 0.2174 -1.000 0.1395 0.03548 0.02245 -0.0504 0.3373 0.2222 -0.750 0.1962 0.03455 0.02123 -0.0562 0.3326 0.2288 -0.500 0.2717 0.03379 0.02000 -0.0658 0.3287 0.2432 -0.250 0.3535 0.03331 0.01932 -0.0770 0.3256 0.2779 0.000 0.4507 0.03212 0.01848 -0.0918 0.3229 0.4146 0.250 0.5034 0.03202 0.01873 -0.0970 0.3218 0.4954 0.500 0.5474 0.03195 0.01919 -0.1005 0.3213 0.5745 0.750 0.6598 0.03292 0.02074 -0.1195 0.3192 1.0001 1.000 0.6911 0.03383 0.02143 -0.1201 0.3180 1.0001 1.250 0.7206 0.03477 0.02223 -0.1203 0.3172 1.0001 1.500 0.7488 0.03575 0.02310 -0.1204 0.3165 1.0001 1.750 0.7759 0.03677 0.02407 -0.1201 0.3165 1.0001 2.000 0.8019 0.03785 0.02519 -0.1197 0.3180 1.0001 2.250 0.8269 0.03903 0.02645 -0.1192 0.3198 1.0001 2.500 0.8513 0.04030 0.02778 -0.1186 0.3221 1.0001 2.750 0.8749 0.04169 0.02923 -0.1179 0.3243 1.0001 3.000 0.8978 0.04320 0.03080 -0.1171 0.3265 1.0001 3.250 0.9209 0.04487 0.03249 -0.1164 0.3287 1.0001 3.500 0.9429 0.04631 0.03408 -0.1155 0.3318 1.0001 3.750 0.9580 0.04775 0.03601 -0.1135 0.3397 1.0001 4.000 0.9763 0.04976 0.03817 -0.1123 0.3449 1.0001 4.250 0.9976 0.05209 0.04051 -0.1116 0.3490 1.0001 4.500 1.0056 0.05389 0.04288 -0.1092 0.3624 1.0001 4.750 1.0254 0.05663 0.04568 -0.1087 0.3691 1.0001 5.000 1.0299 0.05934 0.04888 -0.1067 0.3886 1.0001