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 = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1421 0.06738 0.05893 -0.0184 0.3426 0.4577 -2.750 -0.1181 0.06378 0.05525 -0.0168 0.3397 0.4935 -2.250 -0.0863 0.05740 0.04893 -0.0106 0.3362 0.5947 -2.000 -0.0644 0.05394 0.04546 -0.0093 0.3349 0.6375 -1.750 -0.0341 0.05063 0.04210 -0.0106 0.3332 0.6702 -1.500 0.0201 0.04845 0.03974 -0.0186 0.3316 0.6730 -1.250 0.2577 0.04878 0.03699 -0.0770 0.3283 0.2383 -1.000 0.3081 0.04729 0.03525 -0.0817 0.3277 0.2343 -0.750 0.3546 0.04630 0.03393 -0.0854 0.3276 0.2355 -0.500 0.3982 0.04564 0.03298 -0.0886 0.3269 0.2406 -0.250 0.4391 0.04523 0.03245 -0.0911 0.3255 0.2458 0.000 0.4811 0.04501 0.03191 -0.0939 0.3239 0.2521 0.250 0.5328 0.04511 0.03186 -0.0991 0.3222 0.2640 0.500 0.5921 0.04540 0.03188 -0.1057 0.3215 0.2903 0.750 0.6679 0.04536 0.03212 -0.1163 0.3238 0.3783 1.000 0.7144 0.04601 0.03315 -0.1206 0.3266 0.4444 1.250 0.7507 0.04702 0.03443 -0.1227 0.3293 0.4872 1.500 0.7834 0.04821 0.03590 -0.1241 0.3324 0.5286 1.750 0.8142 0.04942 0.03755 -0.1252 0.3353 0.5841 2.000 0.8449 0.05101 0.03944 -0.1265 0.3381 0.6402 2.250 0.9288 0.05268 0.04220 -0.1398 0.3493 1.0001 2.500 0.9479 0.05484 0.04445 -0.1388 0.3574 1.0001 2.750 0.9702 0.05744 0.04696 -0.1385 0.3633 1.0001 3.000 0.9828 0.05877 0.04860 -0.1364 0.3764 1.0001 3.250 0.9941 0.06179 0.05173 -0.1346 0.3890 1.0001 3.500 0.9909 0.06419 0.05455 -0.1312 0.4116 1.0001 3.750 0.9792 0.06786 0.05860 -0.1277 0.4406 1.0001 4.250 0.2607 0.10807 0.09858 -0.0673 0.8236 0.2793 4.500 0.3163 0.11214 0.10259 -0.0767 0.8077 0.3319 4.750 0.3899 0.11817 0.10910 -0.0889 0.7937 0.4366 5.000 0.4151 0.11996 0.11100 -0.0907 0.7726 0.4769