XFOIL Version 6.96 Calculated polar for: AH 79-100 B AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -4.000 0.4344 0.01396 0.00842 -0.2055 0.8146 0.0341 -3.500 0.4945 0.01257 0.00665 -0.2067 0.8041 0.0354 -3.000 0.5556 0.01067 0.00450 -0.2083 0.7939 0.0383 -2.500 0.6140 0.00998 0.00371 -0.2091 0.7827 0.0414 -2.000 0.6723 0.00959 0.00322 -0.2097 0.7718 0.0460 -1.500 0.7333 0.00908 0.00265 -0.2110 0.7621 0.0661 -1.000 0.7969 0.00818 0.00258 -0.2137 0.7521 0.4032 -0.500 0.8536 0.00809 0.00275 -0.2140 0.7416 0.5474 0.000 0.9092 0.00813 0.00284 -0.2140 0.7302 0.6088 0.500 0.9644 0.00818 0.00296 -0.2138 0.7195 0.6751 1.000 1.0187 0.00820 0.00308 -0.2135 0.7082 0.7263 1.500 1.0721 0.00822 0.00325 -0.2130 0.6966 0.7882 2.000 1.1143 0.00799 0.00328 -0.2097 0.6840 1.0000 2.500 1.1696 0.00819 0.00342 -0.2098 0.6697 1.0000 3.000 1.2234 0.00841 0.00359 -0.2095 0.6538 1.0000 3.500 1.2753 0.00867 0.00377 -0.2088 0.6328 1.0000 4.000 1.3266 0.00895 0.00403 -0.2080 0.6105 1.0000 4.500 1.3771 0.00929 0.00436 -0.2071 0.5885 1.0000 5.000 1.4250 0.00975 0.00474 -0.2056 0.5589 1.0000 5.500 1.4714 0.01028 0.00523 -0.2040 0.5259 1.0000 6.000 1.5143 0.01101 0.00583 -0.2017 0.4848 1.0000 6.500 1.5524 0.01198 0.00661 -0.1985 0.4316 1.0000 7.000 1.5856 0.01322 0.00759 -0.1946 0.3712 1.0000 7.500 1.6094 0.01486 0.00887 -0.1891 0.2951 1.0000 8.000 1.6174 0.01728 0.01074 -0.1811 0.2026 1.0000 8.500 1.6198 0.02025 0.01320 -0.1729 0.1144 1.0000 9.000 1.6259 0.02325 0.01591 -0.1658 0.0627 1.0000 9.500 1.6403 0.02590 0.01854 -0.1605 0.0414 1.0000 10.000 1.6466 0.02946 0.02202 -0.1547 0.0206 1.0000