XFOIL Version 6.94 Calculated polar for: tilt 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 ------- -------- --------- --------- -------- ------- ------- 0.500 0.5929 0.03095 0.01900 -0.0722 0.6215 1.0002 1.000 0.6559 0.03260 0.01916 -0.0720 0.5476 1.0002 1.250 0.6842 0.03371 0.01977 -0.0720 0.5248 1.0002 1.500 0.7134 0.03491 0.02061 -0.0725 0.5087 1.0002 1.750 0.7438 0.03617 0.02163 -0.0734 0.4975 1.0002 2.000 0.7731 0.03756 0.02295 -0.0744 0.4889 1.0002 2.250 0.8018 0.03903 0.02441 -0.0754 0.4814 1.0002 2.500 0.8323 0.04057 0.02584 -0.0765 0.4753 1.0002 2.750 0.8599 0.04238 0.02775 -0.0777 0.4709 1.0002 3.000 0.8844 0.04442 0.03002 -0.0787 0.4673 1.0002 3.250 0.9074 0.04668 0.03247 -0.0796 0.4647 1.0002 3.500 0.9280 0.04922 0.03520 -0.0804 0.4629 1.0002 3.750 0.9458 0.05199 0.03815 -0.0808 0.4607 1.0002 4.000 0.9627 0.05481 0.04109 -0.0811 0.4577 1.0002 4.250 0.9865 0.05718 0.04347 -0.0815 0.4534 1.0002 4.500 0.9917 0.06096 0.04747 -0.0811 0.4512 1.0002 4.750 0.9882 0.06554 0.05229 -0.0803 0.4508 1.0002 5.000 0.9868 0.07033 0.05724 -0.0799 0.4523 1.0002