XFOIL Version 6.94 Calculated polar for: FX S 03-182 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.4344 0.01646 0.01022 -0.1017 0.6974 0.6895 0.500 0.4925 0.01654 0.01025 -0.1034 0.6940 0.6933 1.000 0.5450 0.01684 0.01059 -0.1039 0.6898 0.6967 1.500 0.5983 0.01700 0.01082 -0.1045 0.6868 0.6995 2.000 0.6516 0.01722 0.01112 -0.1051 0.6825 0.7026 2.500 0.7221 0.01564 0.00927 -0.1071 0.6623 0.7056 3.500 0.8237 0.01319 0.00634 -0.1044 0.5771 0.7126 4.500 0.7473 0.02059 0.01085 -0.0764 0.0309 0.7201 5.000 0.7886 0.02141 0.01167 -0.0758 0.0260 0.7239 5.500 0.8238 0.02259 0.01279 -0.0744 0.0108 0.7279 6.000 0.8609 0.02376 0.01400 -0.0733 0.0096 0.7321 6.500 0.8967 0.02512 0.01541 -0.0723 0.0091 0.7368 7.000 0.9325 0.02654 0.01693 -0.0713 0.0089 0.7419 7.500 0.9670 0.02815 0.01862 -0.0705 0.0087 0.7470 8.000 1.0012 0.02981 0.02034 -0.0696 0.0086 0.7521 8.500 1.0330 0.03155 0.02228 -0.0684 0.0088 0.7584 9.000 1.0628 0.03353 0.02446 -0.0670 0.0092 0.7652 9.500 1.0920 0.03561 0.02674 -0.0658 0.0090 0.7731 10.000 1.1190 0.03794 0.02930 -0.0645 0.0089 0.7818 10.500 1.1419 0.04062 0.03231 -0.0627 0.0079 0.7908 11.000 1.1659 0.04338 0.03529 -0.0616 0.0072 0.8010 11.500 1.1818 0.04691 0.03915 -0.0596 0.0067 0.8115 12.000 1.1982 0.05038 0.04293 -0.0579 0.0063 0.8245 12.500 1.2040 0.05512 0.04805 -0.0558 0.0067 0.8382 13.000 1.2137 0.05965 0.05286 -0.0542 0.0067 0.8558 14.000 1.2091 0.07031 0.06427 -0.0488 0.0073 0.9144 14.500 1.2058 0.07493 0.06922 -0.0454 0.0077 1.0000 15.000 1.2143 0.08023 0.07483 -0.0445 0.0079 1.0000 15.500 1.2152 0.08688 0.08191 -0.0432 0.0084 1.0000 16.000 1.2009 0.09614 0.09168 -0.0436 0.0089 1.0000 16.500 1.1692 0.10841 0.10447 -0.0465 0.0093 1.0000 17.000 1.1332 0.12222 0.11875 -0.0523 0.0095 1.0000 17.500 1.1069 0.13545 0.13232 -0.0602 0.0094 1.0000