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 = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1397 0.06705 0.05863 -0.0192 0.3444 0.4439 -2.750 -0.1214 0.06417 0.05574 -0.0174 0.3406 0.4732 -2.500 -0.1052 0.06120 0.05276 -0.0147 0.3379 0.5185 -2.250 -0.0866 0.05802 0.04956 -0.0124 0.3356 0.5646 -2.000 -0.0629 0.05522 0.04669 -0.0120 0.3336 0.6048 -1.750 -0.0291 0.05197 0.04336 -0.0136 0.3314 0.6391 -1.500 0.0209 0.04925 0.04044 -0.0199 0.3294 0.6546 -1.250 0.2638 0.04861 0.03669 -0.0787 0.3247 0.2332 -1.000 0.3167 0.04738 0.03508 -0.0839 0.3239 0.2337 -0.750 0.3654 0.04667 0.03404 -0.0882 0.3233 0.2380 -0.500 0.4090 0.04598 0.03303 -0.0913 0.3233 0.2408 -0.250 0.4496 0.04547 0.03221 -0.0937 0.3237 0.2451 0.000 0.4907 0.04525 0.03200 -0.0964 0.3245 0.2533 0.250 0.5448 0.04522 0.03181 -0.1019 0.3258 0.2697 0.500 0.6005 0.04541 0.03191 -0.1078 0.3272 0.3048 0.750 0.6787 0.04521 0.03203 -0.1189 0.3279 0.4087 1.000 0.7196 0.04591 0.03296 -0.1219 0.3279 0.4663 1.250 0.7545 0.04676 0.03411 -0.1237 0.3280 0.5118 1.500 0.7840 0.04758 0.03540 -0.1243 0.3305 0.5569 1.750 0.8119 0.04849 0.03684 -0.1247 0.3343 0.6221 2.000 0.9067 0.05114 0.04042 -0.1407 0.3412 1.0001 2.250 0.9327 0.05339 0.04255 -0.1408 0.3449 1.0001 2.500 0.9612 0.05637 0.04528 -0.1417 0.3482 1.0001 2.750 0.9664 0.05661 0.04615 -0.1377 0.3621 1.0001 3.000 0.9847 0.05921 0.04873 -0.1367 0.3695 1.0001 3.250 0.9919 0.06099 0.05084 -0.1341 0.3850 1.0001 3.500 1.0076 0.06433 0.05419 -0.1333 0.3970 1.0001 3.750 1.0039 0.06733 0.05757 -0.1303 0.4225 1.0001 4.000 0.8055 0.08157 0.07332 -0.1125 0.5321 1.0001 4.250 0.2851 0.11005 0.10051 -0.0722 0.8261 0.3035 4.500 0.3414 0.11337 0.10412 -0.0823 0.8070 0.3869 4.750 0.3755 0.11594 0.10695 -0.0865 0.7874 0.4496 5.000 0.4040 0.11834 0.10951 -0.0891 0.7673 0.4934