XFOIL Version 6.94 Calculated polar for: lb3000 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2613 0.04784 0.03729 -0.0363 1.0001 0.2519 -2.750 -0.2348 0.04672 0.03588 -0.0372 1.0001 0.2525 -2.500 -0.2084 0.04591 0.03475 -0.0379 1.0001 0.2555 -2.250 -0.1829 0.04521 0.03386 -0.0385 1.0001 0.2604 -2.000 -0.1592 0.04468 0.03332 -0.0386 1.0001 0.2662 -1.750 -0.1344 0.04433 0.03285 -0.0388 1.0001 0.2716 -1.500 -0.1092 0.04414 0.03244 -0.0390 1.0001 0.2773 -1.250 -0.0842 0.04400 0.03224 -0.0392 1.0001 0.2846 -1.000 -0.0595 0.04408 0.03230 -0.0394 1.0001 0.2973 -0.750 -0.0347 0.04418 0.03249 -0.0396 1.0001 0.3136 -0.500 -0.0093 0.04435 0.03279 -0.0400 1.0001 0.3348 0.000 0.0461 0.04306 0.03413 -0.0399 1.0001 0.6697 0.250 0.0775 0.04232 0.03304 -0.0406 0.9890 0.9999 0.500 0.1666 0.04290 0.03298 -0.0525 0.9612 0.9999 0.750 0.2363 0.04301 0.03280 -0.0606 0.9323 0.9999 1.000 0.3107 0.04274 0.03231 -0.0687 0.9021 0.9999 1.250 0.3910 0.04174 0.03117 -0.0769 0.8699 0.9999 1.500 0.4676 0.04028 0.02962 -0.0830 0.8346 0.9999 1.750 0.5318 0.03863 0.02788 -0.0854 0.7976 0.9999 2.000 0.5671 0.03774 0.02691 -0.0828 0.7505 0.9999 2.250 0.5979 0.03655 0.02559 -0.0790 0.7047 0.9999 2.500 0.6208 0.03576 0.02465 -0.0755 0.6547 0.9999 2.750 0.6485 0.03473 0.02341 -0.0728 0.6188 0.9999 3.000 0.6774 0.03389 0.02231 -0.0707 0.5906 0.9999 3.250 0.7065 0.03346 0.02155 -0.0691 0.5676 0.9999 3.500 0.7347 0.03363 0.02144 -0.0681 0.5459 0.9999 3.750 0.7631 0.03403 0.02156 -0.0675 0.5271 0.9999 4.000 0.7905 0.03490 0.02227 -0.0673 0.5095 0.9999 4.250 0.8178 0.03600 0.02327 -0.0675 0.4948 0.9999 4.500 0.8473 0.03686 0.02389 -0.0674 0.4843 0.9999 4.750 0.8732 0.03849 0.02560 -0.0680 0.4733 0.9999