XFOIL Version 6.96 Calculated polar for: FX 61-184 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 ------ -------- --------- --------- -------- -------- -------- -16.000 -0.5999 0.09898 0.09621 -0.0647 1.0000 0.0121 -15.750 -0.6348 0.08793 0.08496 -0.0714 1.0000 0.0121 -15.500 -0.6603 0.07965 0.07651 -0.0761 1.0000 0.0121 -15.250 -0.6855 0.07211 0.06874 -0.0804 1.0000 0.0121 -15.000 -0.7030 0.06637 0.06287 -0.0828 1.0000 0.0120 -14.750 -0.7203 0.06115 0.05744 -0.0852 1.0000 0.0122 -14.500 -0.7409 0.05595 0.05202 -0.0869 1.0000 0.0123 -14.250 -0.7496 0.05251 0.04843 -0.0874 1.0000 0.0123 -14.000 -0.7612 0.04876 0.04455 -0.0875 1.0000 0.0126 -13.750 -0.7366 0.04568 0.04144 -0.0926 0.9540 0.0130 -13.500 -0.7001 0.04237 0.03794 -0.1014 0.9318 0.0134 -13.250 -0.6576 0.03883 0.03410 -0.1112 0.9145 0.0140 -13.000 -0.6168 0.03588 0.03079 -0.1197 0.8931 0.0147 -12.750 -0.5935 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1.4556 0.08052 0.07483 -0.0711 0.0262 0.7816 16.500 1.4598 0.08355 0.07800 -0.0718 0.0252 0.7894 16.750 1.4630 0.08675 0.08132 -0.0726 0.0242 0.7987 17.000 1.4645 0.09025 0.08494 -0.0735 0.0233 0.8105 17.250 1.4632 0.09422 0.08904 -0.0746 0.0224 0.8257 17.500 1.4578 0.09882 0.09379 -0.0759 0.0215 0.8476