XFOIL Version 6.94 Calculated polar for: SC(2)-0714 Supercritical 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 ------- -------- --------- --------- -------- ------- ------- 1.000 0.6505 0.01076 0.00596 -0.1257 0.6754 0.6898 1.500 0.6809 0.01269 0.00652 -0.1207 0.4008 0.6925 2.000 0.7208 0.01424 0.00720 -0.1178 0.2217 0.6950 2.500 0.7688 0.01524 0.00777 -0.1163 0.1371 0.6976 3.000 0.8201 0.01603 0.00833 -0.1154 0.1016 0.7004 3.500 0.8727 0.01670 0.00891 -0.1147 0.0864 0.7040 4.000 0.9273 0.01729 0.00948 -0.1145 0.0778 0.7074 4.500 0.9800 0.01815 0.01026 -0.1141 0.0704 0.7106 5.000 1.0296 0.01874 0.01090 -0.1126 0.0651 0.7134 5.500 1.0733 0.01999 0.01217 -0.1102 0.0602 0.7160 6.000 1.1216 0.02078 0.01306 -0.1086 0.0565 0.7188 6.500 1.1668 0.02195 0.01420 -0.1066 0.0526 0.7219 7.500 1.2587 0.02438 0.01682 -0.1030 0.0472 0.7292 8.000 1.3014 0.02558 0.01804 -0.1008 0.0450 0.7323 10.000 1.4396 0.03183 0.02487 -0.0863 0.0384 0.7448 10.500 1.4709 0.03473 0.02788 -0.0831 0.0369 0.7484 11.000 1.4951 0.03727 0.03072 -0.0787 0.0362 0.7515 11.500 1.5127 0.03983 0.03359 -0.0735 0.0353 0.7545 12.000 1.5255 0.04273 0.03680 -0.0683 0.0343 0.7576 12.500 1.5332 0.04609 0.04043 -0.0633 0.0335 0.7614 13.000 1.5364 0.05004 0.04464 -0.0587 0.0329 0.7654 13.500 1.5358 0.05455 0.04938 -0.0549 0.0324 0.7688 14.000 1.5287 0.06004 0.05513 -0.0519 0.0319 0.7717 14.500 1.5104 0.06729 0.06270 -0.0500 0.0316 0.7742 15.000 1.4751 0.07764 0.07344 -0.0506 0.0313 0.7765 15.500 1.4234 0.09246 0.08873 -0.0564 0.0312 0.7785 16.000 1.3531 0.11429 0.11110 -0.0704 0.0312 0.7797