XFOIL Version 6.94 Calculated polar for: mta3 300.2.14.15.2.c13 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2221 0.04672 0.03574 -0.0268 1.0000 0.3026 -2.750 -0.1834 0.04424 0.03270 -0.0303 1.0000 0.2880 -2.500 -0.1507 0.04189 0.03006 -0.0314 1.0000 0.2809 -2.250 -0.1169 0.03984 0.02768 -0.0324 1.0000 0.2762 -2.000 -0.0837 0.03806 0.02560 -0.0330 1.0000 0.2748 -1.750 -0.0513 0.03647 0.02380 -0.0331 1.0000 0.2763 -1.500 -0.0195 0.03504 0.02224 -0.0328 1.0000 0.2799 -1.250 0.0118 0.03379 0.02094 -0.0322 1.0000 0.2854 -1.000 0.0413 0.03241 0.01991 -0.0312 1.0000 0.2959 -0.750 0.0711 0.03114 0.01914 -0.0301 1.0000 0.3138 -0.500 0.1018 0.02971 0.01860 -0.0289 1.0000 0.3446 -0.250 0.1335 0.02756 0.01831 -0.0279 1.0000 0.4284 0.000 0.1649 0.02572 0.01799 -0.0253 1.0000 1.0000 0.250 0.2867 0.02895 0.01674 -0.0343 0.4737 1.0000 0.500 0.3165 0.03060 0.01732 -0.0334 0.4432 1.0000 0.750 0.3476 0.03218 0.01814 -0.0329 0.4215 1.0000 1.000 0.3803 0.03376 0.01918 -0.0331 0.4059 1.0000 1.250 0.4139 0.03534 0.02044 -0.0335 0.3959 1.0000 1.500 0.4480 0.03707 0.02194 -0.0342 0.3903 1.0000 1.750 0.4819 0.03870 0.02363 -0.0351 0.3872 1.0000 2.000 0.5150 0.04052 0.02556 -0.0362 0.3857 1.0000 2.250 0.5466 0.04252 0.02770 -0.0373 0.3847 1.0000 2.500 0.5766 0.04470 0.03007 -0.0384 0.3837 1.0000 2.750 0.6047 0.04707 0.03264 -0.0396 0.3833 1.0000 3.000 0.6314 0.04969 0.03547 -0.0410 0.3847 1.0000 3.250 0.6572 0.05260 0.03854 -0.0424 0.3886 1.0000 3.500 0.6813 0.05577 0.04208 -0.0454 0.3969 1.0000 3.750 0.7027 0.05951 0.04607 -0.0486 0.4075 1.0000 4.000 0.7190 0.06359 0.05042 -0.0528 0.4207 1.0000 4.250 0.7389 0.06789 0.05479 -0.0558 0.4337 1.0000 4.500 0.7429 0.07283 0.05992 -0.0613 0.4536 1.0000 4.750 0.7456 0.07793 0.06509 -0.0665 0.4753 1.0000 5.000 0.7203 0.08329 0.07050 -0.0728 0.5050 1.0000