XFOIL Version 6.94 Calculated polar for: ARA-D BL 12% 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0465 0.04044 0.03102 -0.0297 0.9997 0.3662 -2.750 -0.0218 0.03943 0.02992 -0.0296 0.9997 0.3838 -2.500 0.0019 0.03850 0.02906 -0.0292 0.9997 0.4046 -2.250 0.0267 0.03775 0.02843 -0.0290 0.9997 0.4312 -2.000 0.0523 0.03717 0.02804 -0.0290 0.9997 0.4656 -1.750 0.0765 0.03656 0.02798 -0.0285 0.9997 0.5119 -1.500 0.0991 0.03601 0.02845 -0.0278 0.9997 0.5856 -1.250 0.2973 0.03104 0.02088 -0.0382 0.5049 1.0003 -1.000 0.3262 0.03249 0.02151 -0.0374 0.4635 1.0003 -0.750 0.3561 0.03391 0.02230 -0.0369 0.4378 1.0003 -0.500 0.3869 0.03533 0.02331 -0.0369 0.4174 1.0003 -0.250 0.4177 0.03680 0.02451 -0.0371 0.4007 1.0003 0.000 0.4481 0.03836 0.02574 -0.0371 0.3880 1.0003 0.250 0.4800 0.04009 0.02750 -0.0379 0.3793 1.0003 0.500 0.5109 0.04193 0.02925 -0.0385 0.3729 1.0003 0.750 0.5411 0.04404 0.03124 -0.0390 0.3682 1.0003 1.000 0.5713 0.04652 0.03395 -0.0406 0.3641 1.0003 1.250 0.5995 0.04917 0.03675 -0.0421 0.3594 1.0003 1.500 0.6262 0.05191 0.03958 -0.0432 0.3551 1.0003 1.750 0.6515 0.05502 0.04277 -0.0445 0.3529 1.0003 2.000 0.6733 0.05920 0.04722 -0.0471 0.3550 1.0003 2.250 0.6890 0.06431 0.05261 -0.0503 0.3597 1.0003 2.500 0.7033 0.06928 0.05770 -0.0528 0.3647 1.0003 2.750 0.7100 0.07506 0.06367 -0.0560 0.3716 1.0003 3.000 0.6927 0.08378 0.07263 -0.0618 0.3875 1.0003 3.500 0.6230 0.10315 0.09224 -0.0754 0.4575 1.0003 4.000 0.4175 0.11761 0.10723 -0.0893 0.8158 1.0003 4.250 0.4404 0.12197 0.11146 -0.0915 0.8038 1.0003 4.500 0.4472 0.12439 0.11376 -0.0909 0.7877 1.0003 5.000 0.4837 0.13196 0.12108 -0.0930 0.7554 1.0003