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 = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4479 0.06507 0.05669 0.0418 0.9999 0.6074 -2.750 -0.4029 0.06352 0.05474 0.0279 0.9999 0.5321 -2.500 -0.3571 0.06181 0.05259 0.0169 0.9999 0.4754 -2.250 -0.3229 0.05942 0.04997 0.0124 0.9999 0.4501 -2.000 -0.2883 0.05748 0.04778 0.0081 0.9999 0.4318 -1.750 -0.2557 0.05552 0.04562 0.0051 0.9999 0.4180 -1.500 -0.2217 0.05367 0.04361 0.0026 0.9999 0.4066 -1.250 -0.1827 0.05209 0.04178 -0.0009 0.9999 0.3954 -1.000 -0.1428 0.05050 0.04003 -0.0038 0.9999 0.3905 -0.750 -0.0995 0.04890 0.03844 -0.0072 0.9999 0.3900 -0.500 0.0663 0.04303 0.03242 -0.0329 0.8318 0.4182 -0.250 0.2545 0.03700 0.02410 -0.0586 0.5773 0.5051 0.000 0.4316 0.03481 0.02154 -0.0891 0.5326 1.0001 0.250 0.4865 0.03616 0.02182 -0.0947 0.5180 1.0001 0.500 0.5307 0.03722 0.02227 -0.0980 0.5066 1.0001 0.750 0.5749 0.03842 0.02290 -0.1013 0.4969 1.0001 1.000 0.6127 0.03936 0.02360 -0.1034 0.4904 1.0001 1.250 0.6510 0.04040 0.02442 -0.1057 0.4861 1.0001 1.500 0.6884 0.04149 0.02537 -0.1079 0.4834 1.0001 1.750 0.7239 0.04265 0.02645 -0.1097 0.4817 1.0001 2.000 0.7571 0.04387 0.02764 -0.1112 0.4807 1.0001 2.250 0.7884 0.04517 0.02898 -0.1124 0.4804 1.0001 2.500 0.8167 0.04653 0.03044 -0.1131 0.4809 1.0001 2.750 0.8419 0.04800 0.03208 -0.1134 0.4819 1.0001 3.000 0.8640 0.04962 0.03387 -0.1133 0.4835 1.0001 3.250 0.8834 0.05140 0.03583 -0.1129 0.4858 1.0001 3.500 0.9005 0.05341 0.03802 -0.1124 0.4886 1.0001 3.750 0.9151 0.05565 0.04044 -0.1116 0.4918 1.0001 4.000 0.9289 0.05811 0.04307 -0.1108 0.4949 1.0001 4.250 0.9431 0.06076 0.04582 -0.1102 0.4977 1.0001 4.500 0.9362 0.06400 0.04931 -0.1074 0.5024 1.0001 4.750 0.9039 0.06882 0.05437 -0.1026 0.5104 1.0001 5.000 0.8940 0.07319 0.05880 -0.1006 0.5166 1.0001