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 = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1933 0.04856 0.04214 0.0109 0.5106 0.7597 -2.750 -0.2035 0.04739 0.04046 0.0146 0.4602 0.7541 -2.500 -0.2118 0.04604 0.03890 0.0164 0.4449 0.7376 -2.250 -0.1950 0.04584 0.03830 0.0106 0.4287 0.6899 -2.000 -0.0217 0.04666 0.03606 -0.0405 0.4047 0.2642 -1.750 0.0074 0.04491 0.03389 -0.0410 0.3972 0.2574 -1.500 0.0406 0.04343 0.03198 -0.0421 0.3902 0.2567 -1.250 0.0817 0.04208 0.03009 -0.0447 0.3831 0.2565 -1.000 0.1313 0.04075 0.02818 -0.0489 0.3762 0.2560 -0.750 0.1781 0.03962 0.02663 -0.0524 0.3689 0.2591 -0.500 0.2424 0.03879 0.02505 -0.0595 0.3612 0.2678 -0.250 0.3253 0.03820 0.02409 -0.0710 0.3561 0.2949 0.000 0.4266 0.03700 0.02270 -0.0862 0.3528 0.3645 0.250 0.5134 0.03561 0.02210 -0.0985 0.3507 0.5268 0.500 0.5664 0.03509 0.02232 -0.1038 0.3498 0.6553 0.750 0.6806 0.03652 0.02392 -0.1229 0.3493 1.0001 1.000 0.7163 0.03755 0.02475 -0.1245 0.3499 1.0001 1.250 0.7481 0.03866 0.02571 -0.1253 0.3508 1.0001 1.500 0.7774 0.03984 0.02678 -0.1256 0.3521 1.0001 1.750 0.8054 0.04111 0.02799 -0.1257 0.3535 1.0001 2.000 0.8321 0.04243 0.02924 -0.1256 0.3546 1.0001 2.250 0.8576 0.04384 0.03061 -0.1253 0.3556 1.0001 2.500 0.8811 0.04530 0.03207 -0.1246 0.3562 1.0001 2.750 0.9041 0.04688 0.03368 -0.1239 0.3571 1.0001 3.000 0.9254 0.04837 0.03525 -0.1229 0.3580 1.0001 3.250 0.9417 0.04934 0.03654 -0.1210 0.3613 1.0001 3.500 0.9573 0.05099 0.03845 -0.1192 0.3659 1.0001 3.750 0.9726 0.05297 0.04062 -0.1176 0.3711 1.0001 4.000 0.9904 0.05517 0.04289 -0.1165 0.3756 1.0001 4.250 1.0128 0.05781 0.04550 -0.1162 0.3792 1.0001 4.500 1.0071 0.05956 0.04790 -0.1121 0.3922 1.0001 4.750 1.0236 0.06240 0.05078 -0.1114 0.3991 1.0001 5.000 1.0094 0.06563 0.05449 -0.1075 0.4160 1.0001