XFOIL Version 6.94 Calculated polar for: coude3 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -2.500 -0.2911 0.07365 0.06688 0.0037 0.9999 0.4053 -2.250 -0.2292 0.06989 0.06272 -0.0091 0.9999 0.3530 -2.000 -0.1833 0.06705 0.05980 -0.0151 0.9999 0.3278 -1.750 -0.1349 0.06468 0.05743 -0.0206 0.9999 0.3069 -1.250 0.1644 0.05270 0.04121 -0.0659 0.4757 0.2655 -1.000 0.2442 0.05173 0.03936 -0.0756 0.4459 0.2692 -0.750 0.3255 0.05126 0.03837 -0.0854 0.4265 0.2737 -0.500 0.4059 0.05133 0.03809 -0.0951 0.4121 0.2817 -0.250 0.4852 0.05172 0.03832 -0.1050 0.3964 0.3032 0.000 0.5470 0.05190 0.03895 -0.1116 0.3850 0.3419 0.250 0.6347 0.05114 0.04001 -0.1222 0.3781 1.0001 0.500 0.7187 0.05388 0.04177 -0.1326 0.3753 1.0001 0.750 0.7869 0.05655 0.04409 -0.1410 0.3749 1.0001 1.000 0.8404 0.05942 0.04678 -0.1467 0.3758 1.0001 1.250 0.8751 0.06202 0.04932 -0.1485 0.3764 1.0001 1.500 0.9011 0.06447 0.05179 -0.1484 0.3765 1.0001 1.750 0.9188 0.06685 0.05422 -0.1467 0.3766 1.0001 2.000 0.9078 0.06805 0.05577 -0.1390 0.3798 1.0001 2.250 0.8887 0.06973 0.05775 -0.1302 0.3843 1.0001 2.500 0.8737 0.07205 0.06026 -0.1226 0.3892 1.0001 2.750 0.8682 0.07494 0.06323 -0.1173 0.3946 1.0001 3.000 0.8496 0.07765 0.06613 -0.1100 0.4023 1.0001 3.250 0.7828 0.08071 0.06962 -0.0958 0.4135 1.0001 3.500 0.7455 0.08479 0.07392 -0.0882 0.4265 1.0001 3.750 0.7305 0.09007 0.07926 -0.0855 0.4438 1.0001 4.000 0.5357 0.10322 0.09363 -0.0766 0.5200 1.0001 4.250 0.1979 0.12357 0.11476 -0.0659 0.9650 0.3239 4.500 0.2076 0.12554 0.11701 -0.0659 0.9667 0.3522 5.000 0.2246 0.12603 0.11893 -0.0645 0.9628 1.0001