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 = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0523 0.06125 0.05233 -0.0186 0.4279 0.8637 -2.750 -0.0108 0.05963 0.05105 -0.0058 0.4317 0.8380 -2.500 -0.0755 0.05760 0.04932 0.0070 0.4354 0.8159 -2.250 -0.1455 0.05532 0.04733 0.0201 0.4401 0.7956 -1.750 -0.0315 0.05813 0.04852 -0.0208 0.4253 0.5028 -1.500 0.0545 0.05689 0.04636 -0.0387 0.4161 0.3788 -1.250 0.1183 0.05567 0.04428 -0.0480 0.4077 0.3427 -1.000 0.1784 0.05376 0.04188 -0.0551 0.4012 0.3340 -0.750 0.2364 0.05257 0.04009 -0.0617 0.3979 0.3217 -0.500 0.2909 0.05122 0.03843 -0.0673 0.3958 0.3197 -0.250 0.3562 0.05038 0.03709 -0.0751 0.3940 0.3224 0.000 0.4238 0.04966 0.03611 -0.0834 0.3927 0.3389 0.250 0.4972 0.04921 0.03532 -0.0929 0.3920 0.3653 0.500 0.5867 0.04838 0.03448 -0.1059 0.3920 0.4259 0.750 0.6630 0.04725 0.03401 -0.1158 0.3927 0.5806 1.000 0.7850 0.04761 0.03553 -0.1370 0.3947 1.0001 1.250 0.8238 0.04934 0.03704 -0.1397 0.3966 1.0001 1.500 0.8591 0.05119 0.03867 -0.1418 0.3986 1.0001 1.750 0.8910 0.05315 0.04043 -0.1431 0.4001 1.0001 2.000 0.9122 0.05440 0.04174 -0.1422 0.4024 1.0001 2.250 0.9284 0.05583 0.04327 -0.1405 0.4050 1.0001 2.500 0.9427 0.05758 0.04511 -0.1387 0.4080 1.0001 2.750 0.9558 0.05958 0.04717 -0.1368 0.4111 1.0001 3.000 0.9652 0.06179 0.04945 -0.1345 0.4145 1.0001 3.250 0.9742 0.06429 0.05198 -0.1322 0.4187 1.0001 3.500 0.9876 0.06703 0.05471 -0.1308 0.4230 1.0001 3.750 0.9878 0.06946 0.05731 -0.1276 0.4302 1.0001 4.000 0.9470 0.07360 0.06186 -0.1195 0.4424 1.0001 4.250 0.9628 0.07722 0.06538 -0.1195 0.4502 1.0001 4.500 0.8523 0.08538 0.07407 -0.1062 0.4742 1.0001