XFOIL Version 6.94 Calculated polar for: GOA 3 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 -0.1501 0.02818 0.01540 -0.0330 0.9999 0.2217 -2.750 -0.1242 0.02492 0.01330 -0.0318 0.9999 0.3453 -2.500 -0.1012 0.02131 0.01161 -0.0272 0.9999 1.0001 -2.250 -0.0758 0.02135 0.01100 -0.0259 0.9999 1.0001 -2.000 -0.0512 0.02141 0.01065 -0.0248 0.9999 1.0001 -1.750 -0.0267 0.02148 0.01037 -0.0237 0.9999 1.0001 -1.500 -0.0022 0.02156 0.01020 -0.0226 0.9999 1.0001 -1.250 0.0226 0.02165 0.01014 -0.0216 0.9999 1.0001 -1.000 0.0479 0.02174 0.01022 -0.0205 0.9999 1.0001 -0.750 0.0746 0.02181 0.01049 -0.0193 0.9999 1.0001 -0.250 0.2048 0.02550 0.01122 -0.0278 0.3238 1.0001 0.000 0.2308 0.02670 0.01198 -0.0268 0.3073 1.0001 0.250 0.2593 0.02789 0.01286 -0.0261 0.2941 1.0001 0.500 0.2893 0.02910 0.01384 -0.0257 0.2796 1.0001 0.750 0.3203 0.03038 0.01493 -0.0256 0.2658 1.0001 1.000 0.3528 0.03183 0.01625 -0.0258 0.2574 1.0001 1.250 0.3871 0.03320 0.01778 -0.0262 0.2562 1.0001 1.500 0.4213 0.03473 0.01950 -0.0266 0.2568 1.0001 1.750 0.4547 0.03643 0.02143 -0.0271 0.2585 1.0001 2.000 0.4873 0.03802 0.02347 -0.0276 0.2626 1.0001 2.250 0.5197 0.03995 0.02584 -0.0283 0.2688 1.0001 2.500 0.5502 0.04230 0.02842 -0.0289 0.2746 1.0001 2.750 0.5819 0.04449 0.03121 -0.0303 0.2855 1.0001 3.000 0.6122 0.04708 0.03422 -0.0317 0.2971 1.0001 3.250 0.6423 0.05001 0.03760 -0.0336 0.3122 1.0001 3.500 0.6720 0.05326 0.04143 -0.0371 0.3347 1.0001 3.750 0.7008 0.05710 0.04564 -0.0404 0.3592 1.0001 4.000 0.7266 0.06150 0.05074 -0.0493 0.4048 1.0001 4.250 0.7434 0.06712 0.05680 -0.0610 0.4646 1.0001 4.500 0.7232 0.07395 0.06372 -0.0749 0.5370 1.0001 4.750 0.6829 0.07942 0.06902 -0.0825 0.6063 1.0001 5.000 0.6355 0.08364 0.07302 -0.0864 0.6887 1.0001