XFOIL Version 6.94 Calculated polar for: NASA/LANGLEY SC(2)-0714 SUPERCR 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 0.5394 0.01074 0.00672 -0.1271 0.8690 0.6818 0.500 0.5955 0.01068 0.00661 -0.1261 0.8235 0.6852 1.000 0.6460 0.01116 0.00643 -0.1240 0.6704 0.6884 1.500 0.6791 0.01317 0.00702 -0.1197 0.3906 0.6922 2.000 0.7260 0.01468 0.00757 -0.1187 0.2048 0.6965 2.500 0.7799 0.01556 0.00800 -0.1187 0.1312 0.6994 3.000 0.8305 0.01620 0.00848 -0.1176 0.1010 0.7022 3.500 0.8808 0.01697 0.00914 -0.1163 0.0855 0.7045 4.000 0.9315 0.01769 0.00981 -0.1152 0.0757 0.7070 4.500 0.9827 0.01838 0.01051 -0.1142 0.0692 0.7097 5.500 1.0822 0.02016 0.01228 -0.1118 0.0592 0.7161 6.000 1.1324 0.02100 0.01314 -0.1107 0.0553 0.7198 7.000 1.2256 0.02297 0.01518 -0.1073 0.0487 0.7253 8.000 1.3091 0.02555 0.01797 -0.1019 0.0439 0.7306 8.500 1.3481 0.02739 0.01980 -0.0992 0.0417 0.7338 9.000 1.3848 0.02857 0.02120 -0.0957 0.0400 0.7373 9.500 1.4158 0.03004 0.02280 -0.0915 0.0384 0.7414 10.000 1.4490 0.03189 0.02468 -0.0880 0.0371 0.7446 10.500 1.4790 0.03434 0.02738 -0.0843 0.0359 0.7478 11.000 1.5027 0.03661 0.02995 -0.0796 0.0348 0.7507 11.500 1.5223 0.03910 0.03271 -0.0747 0.0338 0.7540 12.000 1.5393 0.04172 0.03550 -0.0701 0.0330 0.7576 12.500 1.5546 0.04474 0.03862 -0.0659 0.0321 0.7614 13.000 1.5553 0.04949 0.04370 -0.0610 0.0314 0.7648 13.500 1.5473 0.05462 0.04926 -0.0564 0.0309 0.7681 14.000 1.5297 0.06116 0.05625 -0.0529 0.0305 0.7711 14.500 1.5030 0.06954 0.06507 -0.0514 0.0301 0.7737 15.000 1.4644 0.08085 0.07683 -0.0533 0.0298 0.7760 15.500 1.4052 0.09819 0.09471 -0.0619 0.0298 0.7778