XFOIL Version 6.94 Calculated polar for: nacc632 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 -0.0180 0.02759 0.01406 -0.0242 0.9994 1.0006 -2.750 0.0032 0.02802 0.01425 -0.0239 0.9994 1.0006 -2.500 0.0242 0.02851 0.01452 -0.0235 0.9994 1.0006 -2.250 0.0452 0.02907 0.01489 -0.0231 0.9994 1.0006 -2.000 0.0661 0.02969 0.01534 -0.0228 0.9994 1.0006 -1.750 0.0871 0.03036 0.01587 -0.0225 0.9994 1.0006 -1.500 0.1081 0.03108 0.01645 -0.0223 0.9994 1.0006 -1.250 0.1291 0.03185 0.01712 -0.0221 0.9994 1.0006 -1.000 0.1500 0.03266 0.01785 -0.0220 0.9994 1.0006 -0.750 0.1708 0.03352 0.01866 -0.0220 0.9994 1.0006 -0.500 0.1914 0.03443 0.01953 -0.0219 0.9994 1.0006 -0.250 0.2117 0.03540 0.02049 -0.0220 0.9994 1.0006 0.000 0.2318 0.03644 0.02154 -0.0221 0.9994 1.0006 0.250 0.2514 0.03755 0.02268 -0.0222 0.9994 1.0006 0.500 0.2706 0.03876 0.02395 -0.0225 0.9994 1.0006 0.750 0.2890 0.04009 0.02537 -0.0228 0.9994 1.0006 1.000 0.3064 0.04161 0.02702 -0.0233 0.9994 1.0006 1.250 0.3218 0.04344 0.02900 -0.0241 0.9994 1.0006 1.500 0.3331 0.04591 0.03168 -0.0254 0.9994 1.0006 1.750 0.6118 0.04200 0.02875 -0.0683 0.8201 1.0006 2.000 0.6623 0.04102 0.02779 -0.0656 0.7338 1.0006 2.250 0.6871 0.04137 0.02804 -0.0614 0.6746 1.0006 2.500 0.7121 0.04171 0.02822 -0.0575 0.6303 1.0006 2.750 0.7357 0.04270 0.02919 -0.0552 0.5909 1.0006 3.000 0.7602 0.04365 0.03003 -0.0530 0.5576 1.0006 3.250 0.7865 0.04475 0.03101 -0.0512 0.5320 1.0006 3.500 0.8109 0.04676 0.03311 -0.0511 0.5055 1.0006 3.750 0.8371 0.04805 0.03435 -0.0497 0.4829 1.0006 4.000 0.8627 0.05031 0.03666 -0.0496 0.4645 1.0006 4.250 0.8860 0.05294 0.03943 -0.0499 0.4461 1.0006 4.500 0.9083 0.05565 0.04227 -0.0500 0.4292 1.0006 4.750 0.9290 0.05935 0.04620 -0.0511 0.4179 1.0006 5.000 0.9415 0.06476 0.05195 -0.0534 0.4114 1.0006