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 = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4209 0.06714 0.05732 0.0303 0.9999 0.5598 -2.750 -0.3912 0.06478 0.05472 0.0247 0.9999 0.5340 -2.500 -0.3610 0.06262 0.05232 0.0198 0.9999 0.5138 -2.250 -0.3283 0.06067 0.05005 0.0150 0.9999 0.4948 -2.000 -0.2935 0.05885 0.04795 0.0103 0.9999 0.4770 -1.750 -0.2607 0.05675 0.04566 0.0076 0.9999 0.4660 -1.500 -0.2228 0.05532 0.04387 0.0034 0.9999 0.4554 -1.250 -0.1871 0.05362 0.04202 0.0008 0.9999 0.4525 -1.000 -0.1491 0.05223 0.04042 -0.0022 0.9999 0.4558 -0.750 -0.1096 0.05059 0.03882 -0.0048 0.9999 0.4631 -0.500 -0.0637 0.04923 0.03750 -0.0089 0.9999 0.4711 -0.250 -0.0150 0.04808 0.03648 -0.0139 0.9999 0.4814 0.000 0.1515 0.04281 0.03173 -0.0415 0.8345 0.5666 0.250 0.4126 0.03612 0.02368 -0.0832 0.6407 1.0001 0.750 0.5509 0.03803 0.02269 -0.0981 0.5895 1.0001 1.000 0.5943 0.03926 0.02334 -0.1013 0.5769 1.0001 1.250 0.6357 0.04058 0.02418 -0.1042 0.5657 1.0001 1.500 0.6708 0.04189 0.02525 -0.1061 0.5570 1.0001 1.750 0.7041 0.04327 0.02648 -0.1076 0.5495 1.0001 2.000 0.7392 0.04477 0.02777 -0.1095 0.5438 1.0001 2.250 0.7698 0.04628 0.02922 -0.1108 0.5409 1.0001 2.500 0.7956 0.04784 0.03086 -0.1114 0.5399 1.0001 2.750 0.8191 0.04957 0.03270 -0.1117 0.5396 1.0001 3.000 0.8401 0.05150 0.03476 -0.1118 0.5398 1.0001 3.250 0.8591 0.05359 0.03696 -0.1117 0.5408 1.0001 3.500 0.8736 0.05598 0.03951 -0.1111 0.5423 1.0001 3.750 0.8691 0.05912 0.04293 -0.1087 0.5457 1.0001 4.000 0.8561 0.06313 0.04718 -0.1059 0.5506 1.0001 4.250 0.8398 0.06777 0.05195 -0.1034 0.5565 1.0001 4.500 0.8386 0.07194 0.05614 -0.1026 0.5618 1.0001 4.750 0.7565 0.08079 0.06505 -0.0965 0.5768 1.0001 5.000 0.7625 0.08500 0.06920 -0.0972 0.5842 1.0001