XFOIL Version 6.96 Calculated polar for: FX 61-140 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 ------ -------- --------- --------- -------- -------- -------- -11.750 -0.2950 0.11615 0.11140 -0.0380 1.0000 0.0276 -11.500 -0.2973 0.11187 0.10715 -0.0391 1.0000 0.0261 -11.000 -0.4090 0.10771 0.10266 -0.0410 1.0000 0.0222 -10.750 -0.4084 0.10405 0.09904 -0.0413 1.0000 0.0217 -10.500 -0.4106 0.09934 0.09438 -0.0429 1.0000 0.0214 -10.250 -0.4134 0.09461 0.08970 -0.0447 1.0000 0.0211 -10.000 -0.4185 0.08911 0.08426 -0.0470 1.0000 0.0208 -9.750 -0.4269 0.08268 0.07788 -0.0505 1.0000 0.0204 -9.500 -0.4403 0.07663 0.07184 -0.0539 1.0000 0.0200 -9.250 -0.4572 0.07144 0.06664 -0.0563 1.0000 0.0196 -9.000 -0.4764 0.06721 0.06239 -0.0573 1.0000 0.0194 -8.750 -0.4996 0.06372 0.05887 -0.0567 1.0000 0.0193 -8.500 -0.5219 0.06122 0.05636 -0.0545 1.0000 0.0190 -8.250 -0.5390 0.05805 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-0.0532 0.0361 0.9036 11.250 1.2269 0.04746 0.03996 -0.0527 0.0302 0.9107 11.500 1.2177 0.05114 0.04375 -0.0526 0.0266 0.9197 11.750 1.2103 0.05468 0.04750 -0.0524 0.0235 0.9323 12.000 1.2001 0.05856 0.05154 -0.0525 0.0214 0.9602 12.250 1.1903 0.06333 0.05644 -0.0540 0.0196 1.0000 12.500 1.1844 0.06804 0.06131 -0.0556 0.0181 1.0000 12.750 1.1803 0.07257 0.06603 -0.0571 0.0167 1.0000 13.000 1.1770 0.07706 0.07068 -0.0587 0.0156 1.0000 13.250 1.1743 0.08154 0.07529 -0.0604 0.0147 1.0000 13.500 1.1719 0.08602 0.07989 -0.0621 0.0140 1.0000 13.750 1.1699 0.09047 0.08446 -0.0637 0.0134 1.0000 14.000 1.1683 0.09478 0.08885 -0.0651 0.0128 1.0000