XFOIL Version 6.96 Calculated polar for: FX 61-147 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -10.500 -0.3699 0.09963 0.09475 -0.0501 1.0000 0.0223 -10.250 -0.3730 0.09559 0.09077 -0.0509 1.0000 0.0221 -10.000 -0.3781 0.09110 0.08634 -0.0521 1.0000 0.0219 -9.750 -0.3862 0.08587 0.08118 -0.0540 1.0000 0.0216 -9.500 -0.3971 0.08058 0.07595 -0.0561 1.0000 0.0213 -9.250 -0.4148 0.07543 0.07083 -0.0579 1.0000 0.0210 -9.000 -0.4383 0.07104 0.06647 -0.0586 1.0000 0.0208 -8.750 -0.4625 0.06804 0.06349 -0.0574 1.0000 0.0205 -8.500 -0.4925 0.06587 0.06134 -0.0545 1.0000 0.0205 -8.250 -0.5051 0.06148 0.05682 -0.0576 0.9949 0.0203 -8.000 -0.4973 0.05509 0.05004 -0.0642 0.9839 0.0201 -7.750 -0.4836 0.04957 0.04408 -0.0689 0.9744 0.0200 -7.500 -0.4673 0.04469 0.03871 -0.0718 0.9643 0.0202 -7.250 -0.4469 0.04055 0.03406 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1.0000 13.250 1.2868 0.06591 0.05948 -0.0582 0.0793 1.0000 13.500 1.2774 0.07070 0.06434 -0.0594 0.0705 1.0000 13.750 1.2673 0.07572 0.06941 -0.0608 0.0640 1.0000 14.000 1.2575 0.08086 0.07460 -0.0624 0.0580 1.0000 14.250 1.2491 0.08591 0.07976 -0.0640 0.0523 1.0000 14.500 1.2412 0.09103 0.08494 -0.0658 0.0469 1.0000 14.750 1.2341 0.09614 0.09012 -0.0677 0.0422 1.0000