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 = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4728 0.06681 0.05895 0.0577 0.9999 0.6960 -2.750 -0.4413 0.06582 0.05770 0.0446 0.9999 0.6175 -2.500 -0.3937 0.06419 0.05578 0.0313 0.9999 0.5441 -2.250 -0.3470 0.06265 0.05391 0.0208 0.9999 0.4891 -2.000 -0.3096 0.06054 0.05165 0.0153 0.9999 0.4611 -1.750 -0.2727 0.05858 0.04956 0.0108 0.9999 0.4399 -1.500 -0.2358 0.05670 0.04756 0.0071 0.9999 0.4222 -1.250 -0.1939 0.05499 0.04573 0.0027 0.9999 0.4058 -0.750 0.0179 0.04636 0.03664 -0.0273 0.7770 0.4005 -0.500 0.1909 0.04153 0.02875 -0.0509 0.5611 0.4285 -0.250 0.2720 0.04078 0.02701 -0.0614 0.5354 0.4693 0.000 0.3482 0.03904 0.02525 -0.0707 0.5196 0.5696 0.250 0.5110 0.03804 0.02419 -0.0993 0.5018 1.0001 0.500 0.5638 0.03928 0.02468 -0.1045 0.4943 1.0001 0.750 0.6097 0.04046 0.02532 -0.1083 0.4884 1.0001 1.000 0.6471 0.04146 0.02608 -0.1104 0.4831 1.0001 1.250 0.6838 0.04259 0.02698 -0.1123 0.4779 1.0001 1.500 0.7196 0.04378 0.02800 -0.1142 0.4743 1.0001 1.750 0.7541 0.04508 0.02917 -0.1159 0.4721 1.0001 2.000 0.7861 0.04637 0.03047 -0.1172 0.4719 1.0001 2.250 0.8148 0.04772 0.03188 -0.1179 0.4724 1.0001 2.500 0.8403 0.04920 0.03345 -0.1183 0.4734 1.0001 2.750 0.8626 0.05077 0.03516 -0.1181 0.4751 1.0001 3.000 0.8823 0.05254 0.03708 -0.1177 0.4773 1.0001 3.250 0.8998 0.05450 0.03922 -0.1171 0.4800 1.0001 3.500 0.9149 0.05670 0.04155 -0.1162 0.4832 1.0001 3.750 0.9294 0.05910 0.04407 -0.1154 0.4865 1.0001 4.000 0.9466 0.06167 0.04669 -0.1151 0.4898 1.0001 4.250 0.9280 0.06506 0.05042 -0.1106 0.4963 1.0001 4.500 0.9035 0.06956 0.05513 -0.1064 0.5044 1.0001 4.750 0.9059 0.07340 0.05900 -0.1054 0.5109 1.0001 5.000 0.8107 0.08240 0.06821 -0.0972 0.5292 1.0001