XFOIL Version 6.94 Calculated polar for: FX 61-168 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 0.5271 0.01134 0.00528 -0.1142 0.6014 0.6791 0.500 0.5854 0.01151 0.00541 -0.1150 0.5948 0.6859 1.000 0.6427 0.01148 0.00537 -0.1155 0.5870 0.6910 1.500 0.6985 0.01159 0.00547 -0.1157 0.5795 0.6954 2.000 0.7550 0.01176 0.00567 -0.1161 0.5727 0.7005 2.500 0.8117 0.01181 0.00576 -0.1166 0.5649 0.7056 3.000 0.8697 0.01194 0.00582 -0.1174 0.5562 0.7104 3.500 0.9227 0.01190 0.00592 -0.1170 0.5466 0.7147 4.000 0.9771 0.01197 0.00603 -0.1170 0.5373 0.7187 4.500 1.0309 0.01207 0.00621 -0.1169 0.5260 0.7234 5.000 1.0844 0.01217 0.00630 -0.1167 0.5128 0.7283 5.500 1.1341 0.01212 0.00631 -0.1158 0.4924 0.7322 6.000 1.1800 0.01223 0.00643 -0.1141 0.4659 0.7362 6.500 1.2263 0.01251 0.00675 -0.1126 0.4393 0.7407 7.000 1.2652 0.01309 0.00719 -0.1098 0.3964 0.7450 7.500 1.3028 0.01385 0.00782 -0.1070 0.3600 0.7495 8.000 1.3318 0.01469 0.00855 -0.1027 0.3193 0.7535 8.500 1.3435 0.01611 0.00971 -0.0955 0.2579 0.7577 9.000 1.3460 0.01807 0.01138 -0.0875 0.1960 0.7627 9.500 1.3466 0.02034 0.01343 -0.0801 0.1494 0.7673 10.000 1.3484 0.02289 0.01586 -0.0741 0.1176 0.7712 10.500 1.3418 0.02633 0.01920 -0.0684 0.0835 0.7762 11.000 1.3355 0.03046 0.02328 -0.0641 0.0553 0.7814 11.500 1.3320 0.03503 0.02783 -0.0612 0.0344 0.7862 12.000 1.3310 0.03989 0.03271 -0.0594 0.0210 0.7907 12.500 1.3302 0.04503 0.03793 -0.0581 0.0100 0.7956 13.000 1.3254 0.05092 0.04393 -0.0572 0.0031 0.8007 13.500 1.3293 0.05634 0.04956 -0.0570 0.0025 0.8065 14.500 1.3394 0.06788 0.06156 -0.0581 0.0021 0.8202 15.000 1.3426 0.07441 0.06834 -0.0594 0.0020 0.8280 15.500 1.3439 0.08155 0.07574 -0.0612 0.0020 0.8366 16.000 1.3427 0.08932 0.08380 -0.0635 0.0019 0.8480 16.500 1.3395 0.09767 0.09245 -0.0663 0.0019 0.8634 17.000 1.3311 0.10621 0.10132 -0.0687 0.0019 0.8999