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 = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2818 0.05396 0.04266 -0.0292 1.0001 0.3410 -2.750 -0.2543 0.05233 0.04067 -0.0312 1.0001 0.3434 -2.500 -0.2259 0.05098 0.03892 -0.0330 1.0001 0.3460 -2.250 -0.1990 0.04990 0.03763 -0.0340 1.0001 0.3506 -2.000 -0.1732 0.04918 0.03678 -0.0345 1.0001 0.3601 -1.750 -0.1454 0.04857 0.03592 -0.0356 1.0001 0.3726 -1.500 -0.1192 0.04811 0.03537 -0.0361 1.0001 0.3871 -1.250 -0.0925 0.04768 0.03492 -0.0366 1.0001 0.4027 -1.000 -0.0658 0.04736 0.03463 -0.0370 1.0001 0.4256 -0.750 -0.0400 0.04704 0.03456 -0.0372 1.0001 0.4649 -0.500 -0.0148 0.04640 0.03466 -0.0367 1.0001 0.5400 0.000 0.0200 0.04452 0.03321 -0.0323 1.0001 0.9999 0.250 0.0410 0.04561 0.03386 -0.0326 1.0001 0.9999 0.500 0.0606 0.04681 0.03477 -0.0329 1.0001 0.9999 0.750 0.0789 0.04815 0.03590 -0.0333 1.0001 0.9999 1.000 0.0958 0.04970 0.03731 -0.0338 1.0001 0.9999 1.250 0.1104 0.05155 0.03908 -0.0345 1.0001 0.9999 1.500 0.1215 0.05388 0.04139 -0.0355 0.9999 0.9999 1.750 0.2672 0.05624 0.04341 -0.0596 0.9206 0.9999 2.000 0.3734 0.05735 0.04435 -0.0734 0.8481 0.9999 2.250 0.4442 0.05820 0.04510 -0.0799 0.7893 0.9999 2.500 0.5096 0.05837 0.04515 -0.0842 0.7444 0.9999 2.750 0.5356 0.05982 0.04651 -0.0839 0.7096 0.9999 3.000 0.5790 0.06032 0.04690 -0.0851 0.6845 0.9999 3.250 0.5894 0.06269 0.04920 -0.0839 0.6616 0.9999 3.500 0.6082 0.06451 0.05095 -0.0834 0.6420 0.9999 3.750 0.6395 0.06549 0.05184 -0.0835 0.6254 0.9999 4.000 0.6266 0.07022 0.05656 -0.0822 0.6129 0.9999 4.250 0.6383 0.07320 0.05949 -0.0820 0.6031 0.9999 4.500 0.6237 0.07799 0.06426 -0.0805 0.5954 0.9999 4.750 0.6535 0.07972 0.06592 -0.0812 0.5854 0.9999 5.000 0.6306 0.08527 0.07146 -0.0798 0.5809 0.9999