XFOIL Version 6.94 Calculated polar for: coude3 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3650 0.09935 0.08685 0.0494 0.9999 0.6785 -2.750 -0.3623 0.09560 0.08325 0.0480 0.9999 0.6714 -2.500 -0.3490 0.09196 0.07979 0.0456 0.9999 0.6665 -2.250 -0.3325 0.08846 0.07643 0.0426 0.9999 0.6632 -2.000 -0.3109 0.08519 0.07327 0.0389 0.9999 0.6626 -1.750 -0.2843 0.08220 0.07041 0.0348 0.9999 0.6661 -1.500 -0.2532 0.07951 0.06782 0.0301 0.9999 0.6734 -1.250 -0.2206 0.07713 0.06575 0.0270 0.9999 0.6864 -1.000 -0.1853 0.07498 0.06391 0.0232 0.9999 0.7044 -0.750 -0.1494 0.07308 0.06248 0.0200 0.9999 0.7316 -0.500 -0.1110 0.07133 0.06147 0.0171 0.9999 0.7745 -0.250 -0.0375 0.06929 0.06091 0.0073 0.9999 0.9017 0.000 0.0112 0.06688 0.05917 -0.0054 0.9999 1.0001 0.250 0.0200 0.06943 0.06226 -0.0115 0.9999 1.0001 0.500 -0.0156 0.07358 0.06684 -0.0089 0.9999 1.0001 0.750 -0.0265 0.07715 0.07029 -0.0111 0.9999 1.0001 1.000 -0.0202 0.08047 0.07320 -0.0160 0.9999 1.0001 1.250 -0.0063 0.08372 0.07582 -0.0214 0.9999 1.0001 1.500 0.0101 0.08687 0.07823 -0.0262 0.9999 1.0001 1.750 0.0263 0.08987 0.08045 -0.0298 0.9999 1.0001 2.000 0.0413 0.09272 0.08251 -0.0324 0.9999 1.0001 2.250 0.0552 0.09541 0.08443 -0.0342 0.9999 1.0001 2.500 0.0681 0.09799 0.08628 -0.0355 0.9999 1.0001 2.750 0.0802 0.10049 0.08809 -0.0364 0.9999 1.0001 3.000 0.0919 0.10294 0.08987 -0.0371 0.9999 1.0001 3.250 0.1032 0.10535 0.09165 -0.0377 0.9999 1.0001 3.500 0.1143 0.10774 0.09346 -0.0383 0.9999 1.0001 3.750 0.1253 0.11012 0.09526 -0.0388 0.9999 1.0001 4.000 0.1363 0.11249 0.09712 -0.0394 0.9999 1.0001 4.250 0.1472 0.11486 0.09900 -0.0400 0.9999 1.0001 4.500 0.1581 0.11722 0.10091 -0.0407 0.9999 1.0001 4.750 0.1690 0.11959 0.10285 -0.0413 0.9999 1.0001 5.000 0.1800 0.12197 0.10482 -0.0420 0.9999 1.0001