XFOIL Version 6.94 Calculated polar for: nabad 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 -0.3698 0.05380 0.04125 -0.0029 0.9999 0.1987 -2.750 -0.3385 0.05083 0.03778 -0.0039 0.9999 0.1944 -2.500 -0.3052 0.04816 0.03459 -0.0047 0.9999 0.1922 -2.250 -0.2707 0.04564 0.03165 -0.0054 0.9999 0.1925 -2.000 -0.2348 0.04327 0.02899 -0.0062 0.9999 0.1949 -1.750 -0.1972 0.04112 0.02660 -0.0070 0.9999 0.1992 -1.500 -0.1583 0.03919 0.02445 -0.0080 0.9999 0.2054 -1.250 -0.1198 0.03728 0.02256 -0.0088 0.9999 0.2141 -1.000 -0.0844 0.03554 0.02104 -0.0091 0.9999 0.2284 -0.750 -0.0507 0.03377 0.01975 -0.0090 0.9999 0.2516 -0.500 -0.0185 0.03149 0.01865 -0.0090 0.9999 0.3161 -0.250 0.0602 0.02900 0.01740 -0.0155 0.9999 1.0001 0.000 0.0889 0.02904 0.01728 -0.0146 0.9999 1.0001 0.250 0.1179 0.02913 0.01744 -0.0142 0.9999 1.0001 0.500 0.1465 0.02941 0.01800 -0.0146 0.9999 1.0001 0.750 0.2327 0.03049 0.01886 -0.0258 0.7899 1.0001 1.000 0.2651 0.03165 0.01920 -0.0248 0.7008 1.0001 1.250 0.2922 0.03272 0.01965 -0.0236 0.6616 1.0001 1.500 0.3208 0.03375 0.02024 -0.0234 0.6375 1.0001 1.750 0.3513 0.03480 0.02100 -0.0240 0.6213 1.0001 2.000 0.3815 0.03589 0.02193 -0.0247 0.6109 1.0001 2.250 0.4123 0.03706 0.02308 -0.0261 0.6042 1.0001 2.500 0.4431 0.03830 0.02434 -0.0276 0.6011 1.0001 2.750 0.4742 0.03965 0.02581 -0.0295 0.6007 1.0001 3.000 0.5015 0.04074 0.02696 -0.0299 0.5935 1.0001 3.250 0.5157 0.04047 0.02636 -0.0244 0.5540 1.0001 3.500 0.5260 0.04002 0.02527 -0.0173 0.5061 1.0001 3.750 0.5437 0.04050 0.02542 -0.0139 0.4744 1.0001 4.000 0.5633 0.04117 0.02587 -0.0114 0.4476 1.0001 4.250 0.5832 0.04191 0.02639 -0.0090 0.4196 1.0001 4.500 0.6028 0.04272 0.02702 -0.0067 0.3874 1.0001 4.750 0.6231 0.04372 0.02787 -0.0046 0.3553 1.0001 5.000 0.6440 0.04513 0.02911 -0.0025 0.3231 1.0001