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 = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -14.750 -0.4647 0.08409 0.08159 -0.0858 0.9469 0.0153 -14.500 -0.3926 0.08427 0.08178 -0.0953 0.9217 0.0159 -14.250 -0.4857 0.06180 0.05859 -0.1126 0.9137 0.0145 -14.000 -0.4939 0.05453 0.05089 -0.1218 0.8837 0.0146 -13.750 -0.5036 0.05027 0.04629 -0.1248 0.8560 0.0146 -13.500 -0.5192 0.04636 0.04206 -0.1260 0.8340 0.0147 -13.250 -0.5245 0.04418 0.03965 -0.1264 0.8155 0.0148 -13.000 -0.5334 0.04109 0.03632 -0.1262 0.8001 0.0147 -12.750 -0.5478 0.03751 0.03244 -0.1257 0.7860 0.0150 -12.500 -0.5481 0.03521 0.02995 -0.1249 0.7726 0.0152 -12.250 -0.5437 0.03348 0.02808 -0.1241 0.7605 0.0154 -12.000 -0.5378 0.03211 0.02659 -0.1234 0.7485 0.0157 -11.500 -0.5236 0.02947 0.02367 -0.1216 0.7272 0.0162 -11.250 -0.5142 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0.00573 -0.1116 0.5356 0.6024 3.500 0.8925 0.01267 0.00593 -0.1119 0.5346 0.6056 3.750 0.9195 0.01275 0.00609 -0.1120 0.5335 0.6090 4.000 0.9468 0.01286 0.00626 -0.1122 0.5324 0.6127 4.250 0.9743 0.01300 0.00645 -0.1125 0.5314 0.6165 4.500 1.0016 0.01313 0.00665 -0.1127 0.5304 0.6203 4.750 1.0288 0.01326 0.00686 -0.1130 0.5294 0.6245 5.000 1.0562 0.01340 0.00707 -0.1132 0.5283 0.6292 5.250 1.0837 0.01355 0.00726 -0.1135 0.5272 0.6342 5.500 1.1109 0.01368 0.00749 -0.1138 0.5262 0.6392 5.750 1.1385 0.01384 0.00772 -0.1141 0.5253 0.6451 6.000 1.1662 0.01399 0.00793 -0.1144 0.5242 0.6513 6.250 1.1942 0.01412 0.00813 -0.1148 0.5230 0.6580 6.500 1.2229 0.01431 0.00836 -0.1153 0.5217 0.6657 6.750 1.2512 0.01463 0.00875 -0.1158 0.5200 0.6735 7.000 1.2753 0.01473 0.00898 -0.1155 0.5187 0.6826 7.250 1.2989 0.01483 0.00924 -0.1151 0.5168 0.6923 7.500 1.3229 0.01494 0.00949 -0.1147 0.5148 0.7031 7.750 1.3475 0.01504 0.00971 -0.1145 0.5127 0.7156 8.000 1.3740 0.01491 0.00965 -0.1144 0.5093 0.7298 8.250 1.3963 0.01455 0.00935 -0.1134 0.5016 0.7461 8.500 1.4141 0.01399 0.00883 -0.1113 0.4918 0.7654 8.750 1.4261 0.01376 0.00877 -0.1085 0.4829 0.7897 9.000 1.4417 0.01357 0.00867 -0.1062 0.4755 0.8202 9.250 1.4527 0.01353 0.00887 -0.1032 0.4697 0.8661 9.500 1.4751 0.01340 0.00896 -0.1026 0.4620 1.0000 9.750 1.4871 0.01358 0.00920 -0.1001 0.4544 1.0000 10.000 1.4974 0.01383 0.00945 -0.0974 0.4455 1.0000 10.250 1.5056 0.01422 0.00990 -0.0946 0.4320 1.0000 10.500 1.5089 0.01486 0.01054 -0.0913 0.4061 1.0000 10.750 1.4904 0.01663 0.01210 -0.0856 0.3690 1.0000 11.000 1.4666 0.01938 0.01467 -0.0805 0.3328 1.0000 11.250 1.4446 0.02262 0.01778 -0.0767 0.3041 1.0000 11.500 1.4198 0.02652 0.02156 -0.0734 0.2734 1.0000 11.750 1.3967 0.03062 0.02557 -0.0708 0.2488 1.0000 12.000 1.3750 0.03482 0.02968 -0.0686 0.2247 1.0000 12.250 1.3512 0.03942 0.03417 -0.0666 0.1994 1.0000 12.500 1.3311 0.04382 0.03847 -0.0650 0.1737 1.0000 12.750 1.3137 0.04808 0.04261 -0.0637 0.1500 1.0000 13.000 1.2969 0.05236 0.04676 -0.0625 0.1242 1.0000 13.250 1.2821 0.05649 0.05075 -0.0614 0.0987 1.0000 13.500 1.2676 0.06069 0.05476 -0.0605 0.0718 1.0000 13.750 1.2589 0.06432 0.05827 -0.0598 0.0522 1.0000 14.000 1.2501 0.06805 0.06189 -0.0591 0.0364 1.0000 14.250 1.2483 0.07111 0.06491 -0.0587 0.0283 1.0000 14.500 1.2463 0.07427 0.06805 -0.0584 0.0234 1.0000 14.750 1.2498 0.07681 0.07064 -0.0582 0.0211 1.0000 15.000 1.2502 0.07973 0.07358 -0.0580 0.0191 1.0000 15.250 1.2516 0.08260 0.07651 -0.0579 0.0176 1.0000 15.500 1.2555 0.08520 0.07916 -0.0579 0.0167 1.0000 15.750 1.2579 0.08800 0.08202 -0.0579 0.0158 1.0000 16.000 1.2579 0.09109 0.08515 -0.0579 0.0150 1.0000