XFOIL Version 6.94 Calculated polar for: mta2 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2893 0.04942 0.04011 -0.0223 1.0000 0.2955 -2.750 -0.2470 0.04648 0.03669 -0.0269 1.0000 0.2746 -2.500 -0.2100 0.04390 0.03381 -0.0290 1.0000 0.2607 -2.250 -0.1715 0.04175 0.03122 -0.0308 1.0000 0.2484 -2.000 -0.1332 0.03992 0.02899 -0.0319 1.0000 0.2379 -1.750 -0.0975 0.03823 0.02717 -0.0321 1.0000 0.2325 -1.500 -0.0611 0.03666 0.02566 -0.0319 1.0000 0.2291 -1.250 -0.0230 0.03522 0.02443 -0.0316 1.0000 0.2295 -1.000 0.0051 0.03319 0.02287 -0.0297 0.7399 0.2331 -0.750 0.1390 0.03430 0.02053 -0.0454 0.3549 0.2623 -0.500 0.1747 0.03399 0.02014 -0.0456 0.3407 0.2912 0.000 0.2520 0.03221 0.01971 -0.0449 0.3243 1.0000 0.250 0.2948 0.03332 0.01992 -0.0456 0.3194 1.0000 0.500 0.3367 0.03446 0.02046 -0.0469 0.3160 1.0000 0.750 0.3781 0.03573 0.02131 -0.0485 0.3141 1.0000 1.000 0.4180 0.03707 0.02244 -0.0501 0.3139 1.0000 1.250 0.4561 0.03851 0.02381 -0.0516 0.3151 1.0000 1.500 0.4924 0.04007 0.02537 -0.0530 0.3174 1.0000 1.750 0.5270 0.04179 0.02713 -0.0542 0.3205 1.0000 2.000 0.5600 0.04353 0.02901 -0.0552 0.3247 1.0000 2.250 0.5909 0.04530 0.03111 -0.0563 0.3320 1.0000 2.500 0.6206 0.04749 0.03340 -0.0572 0.3380 1.0000 2.750 0.6476 0.04955 0.03575 -0.0581 0.3454 1.0000 3.000 0.6725 0.05196 0.03837 -0.0591 0.3534 1.0000 3.250 0.6971 0.05451 0.04107 -0.0599 0.3607 1.0000 3.500 0.7177 0.05723 0.04406 -0.0613 0.3720 1.0000 3.750 0.7363 0.06010 0.04721 -0.0631 0.3852 1.0000 4.000 0.7524 0.06331 0.05064 -0.0651 0.4009 1.0000 4.250 0.7630 0.06692 0.05450 -0.0678 0.4217 1.0000 4.500 0.7753 0.07098 0.05868 -0.0712 0.4460 1.0000 4.750 0.7698 0.07557 0.06345 -0.0761 0.4815 1.0000 5.000 0.7346 0.08073 0.06868 -0.0816 0.5288 1.0000