XFOIL Version 6.94 Calculated polar for: rt 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2901 0.04570 0.03408 0.0509 0.9991 0.6628 -2.750 -0.2649 0.04637 0.03481 0.0497 0.9991 0.6725 -2.500 -0.2399 0.04711 0.03557 0.0484 0.9991 0.6822 -2.250 -0.2140 0.04791 0.03640 0.0467 0.9991 0.6923 -2.000 -0.1171 0.04904 0.03772 0.0315 0.9725 0.7189 -1.750 0.0139 0.04799 0.03695 0.0117 0.9186 0.7516 -1.500 0.2040 0.04347 0.03339 -0.0158 0.8526 0.8359 -1.250 0.5210 0.03220 0.02221 -0.0577 0.6432 0.9600 -1.000 0.5554 0.03321 0.02182 -0.0606 0.5636 1.0009 -0.750 0.5825 0.03388 0.02174 -0.0608 0.5210 1.0009 -0.500 0.6106 0.03461 0.02198 -0.0612 0.4946 1.0009 -0.250 0.6405 0.03535 0.02229 -0.0619 0.4753 1.0009 0.000 0.6725 0.03603 0.02260 -0.0630 0.4618 1.0009 0.250 0.7018 0.03694 0.02340 -0.0638 0.4503 1.0009 0.500 0.7345 0.03769 0.02390 -0.0651 0.4419 1.0009 0.750 0.7632 0.03872 0.02489 -0.0657 0.4337 1.0009 1.000 0.7904 0.03973 0.02587 -0.0661 0.4271 1.0009 1.250 0.8199 0.04068 0.02667 -0.0668 0.4214 1.0009 1.500 0.8470 0.04185 0.02777 -0.0672 0.4168 1.0009 1.750 0.8668 0.04336 0.02945 -0.0664 0.4124 1.0009 2.000 0.8873 0.04495 0.03115 -0.0658 0.4093 1.0009 2.250 0.9067 0.04670 0.03302 -0.0651 0.4073 1.0009 2.500 0.9240 0.04867 0.03512 -0.0642 0.4057 1.0009 2.750 0.9380 0.05094 0.03756 -0.0629 0.4049 1.0009 3.000 0.9472 0.05357 0.04037 -0.0611 0.4044 1.0009 3.250 0.9480 0.05690 0.04393 -0.0585 0.4048 1.0009 3.500 0.9354 0.06135 0.04865 -0.0546 0.4060 1.0009 3.750 0.8993 0.06788 0.05543 -0.0489 0.4090 1.0009 4.000 0.8361 0.07711 0.06484 -0.0423 0.4137 1.0009 4.250 0.7806 0.08615 0.07393 -0.0379 0.4198 1.0009