XFOIL Version 6.96 Calculated polar for: FX 66-S-196 V1 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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -17.000 -0.5970 0.10336 0.10057 -0.0640 0.9570 0.0082 -16.750 -0.6207 0.09100 0.08795 -0.0741 0.9462 0.0083 -16.500 -0.6282 0.08075 0.07744 -0.0849 0.9334 0.0082 -16.250 -0.6229 0.07137 0.06773 -0.0974 0.9134 0.0084 -16.000 -0.6132 0.06434 0.06032 -0.1074 0.8842 0.0085 -15.750 -0.6127 0.05962 0.05529 -0.1119 0.8563 0.0087 -15.500 -0.6159 0.05586 0.05130 -0.1141 0.8339 0.0086 -15.250 -0.6148 0.05298 0.04820 -0.1156 0.8146 0.0089 -15.000 -0.6173 0.04997 0.04499 -0.1166 0.7976 0.0088 -14.750 -0.6181 0.04732 0.04222 -0.1176 0.7833 0.0091 -14.500 -0.6187 0.04484 0.03958 -0.1180 0.7671 0.0092 -14.000 -0.6137 0.04076 0.03523 -0.1186 0.7418 0.0094 -13.750 -0.6096 0.03895 0.03330 -0.1187 0.7301 0.0095 -13.500 -0.6044 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-0.1099 0.4948 0.5916 6.000 1.1484 0.01230 0.00580 -0.1100 0.4937 0.5964 6.250 1.1752 0.01243 0.00598 -0.1101 0.4926 0.6012 6.500 1.2020 0.01256 0.00618 -0.1103 0.4916 0.6062 6.750 1.2289 0.01271 0.00638 -0.1104 0.4901 0.6121 7.250 1.2795 0.01297 0.00684 -0.1102 0.4873 0.6242 7.500 1.3033 0.01307 0.00707 -0.1098 0.4851 0.6312 7.750 1.3268 0.01316 0.00727 -0.1093 0.4825 0.6382 8.000 1.3501 0.01325 0.00746 -0.1088 0.4798 0.6461 8.250 1.3693 0.01325 0.00749 -0.1074 0.4732 0.6545 8.500 1.3810 0.01322 0.00756 -0.1046 0.4618 0.6642 8.750 1.3869 0.01330 0.00769 -0.1007 0.4477 0.6746 9.000 1.3926 0.01357 0.00800 -0.0970 0.4314 0.6861 9.500 1.3779 0.01532 0.00958 -0.0864 0.3666 0.7121 9.750 1.3581 0.01715 0.01130 -0.0804 0.3337 0.7279 10.250 1.3241 0.02169 0.01579 -0.0712 0.2819 0.7700 10.500 1.3039 0.02485 0.01894 -0.0677 0.2545 0.8011 10.750 1.2929 0.02797 0.02215 -0.0659 0.2298 0.8541 11.250 1.2722 0.03647 0.03058 -0.0665 0.1705 1.0000 11.500 1.2593 0.04011 0.03414 -0.0651 0.1490 1.0000 11.750 1.2484 0.04369 0.03764 -0.0639 0.1285 1.0000 12.000 1.2406 0.04702 0.04090 -0.0629 0.1093 1.0000 12.250 1.2332 0.05038 0.04417 -0.0620 0.0902 1.0000 12.500 1.2273 0.05363 0.04733 -0.0612 0.0718 1.0000 12.750 1.2249 0.05659 0.05023 -0.0606 0.0581 1.0000 13.000 1.2234 0.05949 0.05307 -0.0600 0.0458 1.0000 13.250 1.2237 0.06226 0.05580 -0.0596 0.0365 1.0000 13.500 1.2255 0.06486 0.05838 -0.0593 0.0287 1.0000 13.750 1.2284 0.06741 0.06093 -0.0590 0.0235 1.0000 14.000 1.2329 0.06979 0.06332 -0.0588 0.0196 1.0000 14.250 1.2377 0.07218 0.06573 -0.0586 0.0171 1.0000 14.500 1.2419 0.07465 0.06822 -0.0585 0.0149 1.0000 14.750 1.2478 0.07691 0.07053 -0.0584 0.0138 1.0000 15.000 1.2530 0.07930 0.07295 -0.0583 0.0125 1.0000 15.250 1.2571 0.08183 0.07551 -0.0583 0.0115 1.0000 15.500 1.2625 0.08423 0.07796 -0.0583 0.0109 1.0000 15.750 1.2681 0.08663 0.08043 -0.0583 0.0103 1.0000 16.000 1.2729 0.08914 0.08300 -0.0584 0.0098 1.0000 16.250 1.2774 0.09166 0.08557 -0.0585 0.0093 1.0000 16.500 1.2805 0.09440 0.08836 -0.0587 0.0088 1.0000