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 = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2079 0.04757 0.03907 -0.0097 1.0000 0.3801 -2.750 -0.1817 0.04525 0.03647 -0.0117 1.0000 0.3889 -2.500 -0.1597 0.04343 0.03465 -0.0115 1.0000 0.4029 -2.250 -0.1335 0.04171 0.03277 -0.0124 1.0000 0.4190 -2.000 -0.1069 0.04032 0.03125 -0.0132 1.0000 0.4400 -1.750 -0.0820 0.03909 0.03007 -0.0133 1.0000 0.4659 -1.500 -0.0584 0.03798 0.02922 -0.0128 1.0000 0.4976 -1.250 -0.0336 0.03707 0.02862 -0.0126 1.0000 0.5404 -1.000 -0.0122 0.03626 0.02847 -0.0118 1.0000 0.5972 -0.750 0.0007 0.03581 0.02914 -0.0101 1.0000 0.6768 -0.500 0.3046 0.03374 0.02365 -0.0369 0.5186 1.0000 -0.250 0.3315 0.03514 0.02448 -0.0359 0.4907 1.0000 0.000 0.3593 0.03672 0.02568 -0.0356 0.4711 1.0000 0.250 0.3867 0.03831 0.02690 -0.0352 0.4553 1.0000 0.500 0.4135 0.03998 0.02832 -0.0350 0.4407 1.0000 0.750 0.4393 0.04200 0.03025 -0.0353 0.4288 1.0000 1.000 0.4650 0.04399 0.03210 -0.0354 0.4193 1.0000 1.250 0.4913 0.04626 0.03420 -0.0357 0.4138 1.0000 1.500 0.5142 0.04948 0.03760 -0.0376 0.4111 1.0000 1.750 0.5338 0.05307 0.04132 -0.0396 0.4099 1.0000 2.000 0.5490 0.05701 0.04536 -0.0415 0.4092 1.0000 2.250 0.5599 0.06127 0.04968 -0.0435 0.4093 1.0000 2.500 0.5674 0.06574 0.05418 -0.0453 0.4102 1.0000 2.750 0.5736 0.07027 0.05871 -0.0471 0.4120 1.0000 3.250 0.5458 0.08285 0.07139 -0.0545 0.4335 1.0000 3.500 0.5161 0.08946 0.07801 -0.0586 0.4544 1.0000 3.750 0.5015 0.09523 0.08372 -0.0620 0.4795 1.0000 4.000 0.4550 0.10173 0.09030 -0.0677 0.5448 1.0000 4.500 0.3007 0.10614 0.09506 -0.0711 0.8996 1.0000 4.750 0.3135 0.10950 0.09825 -0.0721 0.8957 1.0000 5.000 0.3229 0.11166 0.10030 -0.0721 0.8818 1.0000