XFOIL Version 6.94 Calculated polar for: FX 61-147 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.5462 0.01014 0.00417 -0.1195 0.5738 0.7010 1.000 0.6577 0.01055 0.00436 -0.1198 0.5425 0.7147 1.500 0.7110 0.01072 0.00452 -0.1194 0.5287 0.7198 2.000 0.7647 0.01099 0.00470 -0.1191 0.5142 0.7255 2.500 0.8203 0.01116 0.00486 -0.1193 0.5009 0.7318 3.000 0.8749 0.01140 0.00507 -0.1193 0.4887 0.7371 3.500 0.9279 0.01155 0.00529 -0.1189 0.4764 0.7415 4.000 0.9814 0.01185 0.00559 -0.1187 0.4646 0.7469 4.500 1.0356 0.01210 0.00588 -0.1187 0.4529 0.7524 5.000 1.0840 0.01228 0.00599 -0.1175 0.4269 0.7569 5.500 1.1337 0.01249 0.00628 -0.1165 0.4094 0.7617 6.000 1.1810 0.01283 0.00659 -0.1152 0.3843 0.7666 6.500 1.2265 0.01329 0.00696 -0.1137 0.3468 0.7716 7.000 1.2709 0.01388 0.00747 -0.1120 0.3139 0.7764 7.500 1.3086 0.01469 0.00815 -0.1092 0.2691 0.7808 8.000 1.3210 0.01662 0.00951 -0.1022 0.1714 0.7861 8.500 1.3278 0.01897 0.01140 -0.0948 0.0931 0.7916 9.000 1.3411 0.02094 0.01318 -0.0888 0.0532 0.7961 9.500 1.3484 0.02315 0.01531 -0.0823 0.0236 0.8017 10.000 1.3614 0.02526 0.01750 -0.0773 0.0126 0.8074 10.500 1.3634 0.02837 0.02068 -0.0720 0.0022 0.8127 11.000 1.3738 0.03127 0.02378 -0.0685 0.0017 0.8180 11.500 1.3805 0.03476 0.02751 -0.0654 0.0015 0.8245 12.000 1.3843 0.03906 0.03205 -0.0632 0.0014 0.8313 12.500 1.3850 0.04416 0.03742 -0.0619 0.0014 0.8376 13.000 1.3826 0.05009 0.04364 -0.0614 0.0013 0.8449 13.500 1.3778 0.05693 0.05075 -0.0618 0.0013 0.8528 14.000 1.3715 0.06438 0.05850 -0.0630 0.0013 0.8620 15.000 1.3493 0.08111 0.07589 -0.0662 0.0013 0.9109