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 = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2711 0.05146 0.04029 -0.0332 1.0001 0.3052 -2.750 -0.2473 0.05026 0.03898 -0.0337 1.0001 0.3101 -2.500 -0.2211 0.04917 0.03765 -0.0347 1.0001 0.3145 -2.250 -0.1938 0.04824 0.03641 -0.0358 1.0001 0.3182 -2.000 -0.1659 0.04755 0.03535 -0.0368 1.0001 0.3232 -1.750 -0.1402 0.04699 0.03471 -0.0372 1.0001 0.3314 -1.500 -0.1138 0.04672 0.03424 -0.0377 1.0001 0.3446 -1.250 -0.0884 0.04643 0.03399 -0.0380 1.0001 0.3594 -1.000 -0.0621 0.04625 0.03378 -0.0384 1.0001 0.3768 -0.750 -0.0358 0.04609 0.03375 -0.0388 1.0001 0.4015 -0.500 -0.0097 0.04587 0.03394 -0.0390 1.0001 0.4487 -0.250 0.0141 0.04443 0.03430 -0.0371 1.0001 0.6408 0.000 0.0216 0.04335 0.03309 -0.0332 1.0001 0.9999 0.250 0.0432 0.04447 0.03368 -0.0335 1.0001 0.9999 0.500 0.0629 0.04570 0.03461 -0.0338 1.0001 0.9999 0.750 0.0812 0.04709 0.03579 -0.0342 1.0001 0.9999 1.000 0.0977 0.04871 0.03729 -0.0348 1.0001 0.9999 1.250 0.1747 0.05038 0.03868 -0.0466 0.9700 0.9999 1.500 0.2838 0.05131 0.03934 -0.0622 0.9137 0.9999 1.750 0.3824 0.05128 0.03916 -0.0738 0.8526 0.9999 2.000 0.4788 0.05029 0.03808 -0.0828 0.7933 0.9999 2.250 0.5439 0.04930 0.03696 -0.0853 0.7441 0.9999 2.500 0.5831 0.04906 0.03659 -0.0845 0.7043 0.9999 2.750 0.6055 0.04987 0.03731 -0.0829 0.6696 0.9999 3.000 0.6434 0.04923 0.03652 -0.0820 0.6458 0.9999 3.250 0.6597 0.05070 0.03793 -0.0808 0.6201 0.9999 3.500 0.6877 0.05118 0.03830 -0.0800 0.6015 0.9999 3.750 0.7237 0.05087 0.03785 -0.0794 0.5877 0.9999 4.000 0.7367 0.05352 0.04051 -0.0792 0.5718 0.9999 4.250 0.7506 0.05609 0.04305 -0.0788 0.5584 0.9999 4.500 0.7936 0.05493 0.04173 -0.0782 0.5479 0.9999 4.750 0.7787 0.06131 0.04823 -0.0782 0.5361 0.9999 5.000 0.8154 0.06144 0.04828 -0.0781 0.5284 0.9999