XFOIL Version 6.94 Calculated polar for: AH 82-150 A 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 0.4278 0.01154 0.00659 -0.0944 0.7180 0.7928 0.500 0.4853 0.01137 0.00639 -0.0953 0.7111 0.7959 1.000 0.5440 0.01124 0.00620 -0.0965 0.7041 0.7992 1.500 0.6037 0.01113 0.00601 -0.0979 0.6975 0.8011 2.000 0.6626 0.01125 0.00609 -0.0993 0.6901 0.8027 2.500 0.7198 0.01117 0.00609 -0.1003 0.6836 0.8041 3.000 0.7766 0.01093 0.00588 -0.1010 0.6753 0.8072 3.500 0.8341 0.01081 0.00578 -0.1019 0.6673 0.8094 4.000 0.8907 0.01088 0.00589 -0.1026 0.6580 0.8115 4.500 0.9456 0.01080 0.00599 -0.1029 0.6480 0.8138 5.000 1.0020 0.01070 0.00593 -0.1034 0.6354 0.8163 5.500 1.0535 0.01050 0.00575 -0.1028 0.6070 0.8189 6.000 1.0982 0.01063 0.00572 -0.1010 0.5505 0.8218 6.500 1.1297 0.01156 0.00630 -0.0971 0.4663 0.8249 7.000 1.1331 0.01362 0.00766 -0.0888 0.3414 0.8282 7.500 1.0398 0.02148 0.01350 -0.0694 0.0170 0.8351 8.000 1.0696 0.02316 0.01538 -0.0671 0.0160 0.8388 8.500 1.0914 0.02549 0.01794 -0.0641 0.0147 0.8430 9.000 1.1143 0.02783 0.02043 -0.0616 0.0142 0.8474 9.500 1.1328 0.03064 0.02338 -0.0590 0.0127 0.8519 10.000 1.1471 0.03375 0.02670 -0.0563 0.0116 0.8564 10.500 1.1542 0.03757 0.03073 -0.0530 0.0114 0.8619 11.000 1.1615 0.04158 0.03493 -0.0501 0.0106 0.8681 11.500 1.1775 0.04506 0.03859 -0.0478 0.0110 0.8744 12.000 1.1894 0.04912 0.04289 -0.0435 0.0095 0.8804 12.500 1.2086 0.05229 0.04630 -0.0423 0.0088 0.8907 13.000 1.2240 0.05607 0.05036 -0.0405 0.0074 0.9046 13.500 1.2328 0.05998 0.05469 -0.0364 0.0069 0.9480 14.000 1.2404 0.06577 0.06090 -0.0342 0.0069 1.0000 14.500 1.2317 0.07367 0.06927 -0.0330 0.0072 1.0000 15.000 1.2069 0.08401 0.08012 -0.0331 0.0076 1.0000 15.500 1.1754 0.09574 0.09228 -0.0356 0.0080 1.0000 16.000 1.1379 0.10927 0.10621 -0.0407 0.0084 1.0000 16.500 1.0961 0.12502 0.12232 -0.0488 0.0087 1.0000