XFOIL Version 6.96 Calculated polar for: BOEING VERTOL V43012-1.58 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -3.000 -0.0174 0.01938 0.01325 -0.0330 0.6988 0.0528 -2.500 0.0293 0.01796 0.01137 -0.0304 0.6505 0.0596 -2.000 0.0784 0.01735 0.01041 -0.0281 0.6043 0.0656 -1.500 0.1282 0.01528 0.00808 -0.0264 0.5601 0.0695 -0.500 0.2250 0.01149 0.00350 -0.0206 0.4395 0.0384 0.000 0.2691 0.01134 0.00307 -0.0180 0.3883 0.0394 0.500 0.3173 0.01137 0.00292 -0.0163 0.3412 0.0422 1.000 0.3657 0.01146 0.00280 -0.0145 0.3043 0.0452 1.500 0.4153 0.01159 0.00283 -0.0130 0.2810 0.0589 2.000 0.4469 0.01024 0.00284 -0.0087 0.2665 0.5975 2.500 0.5113 0.00984 0.00345 -0.0097 0.2530 0.9450 3.000 0.6087 0.01050 0.00397 -0.0181 0.2406 0.9752 3.500 0.7188 0.01108 0.00444 -0.0294 0.2305 0.9916 4.000 0.8111 0.01151 0.00472 -0.0374 0.2213 1.0000 4.500 0.8526 0.01169 0.00492 -0.0347 0.2167 1.0000 5.000 0.8929 0.01195 0.00516 -0.0318 0.2111 1.0000 5.500 0.9312 0.01239 0.00555 -0.0286 0.2054 1.0000 6.000 0.9719 0.01267 0.00589 -0.0257 0.2004 1.0000 6.500 1.0108 0.01309 0.00628 -0.0227 0.1947 1.0000 7.000 1.0494 0.01360 0.00681 -0.0195 0.1889 1.0000 7.500 1.0899 0.01396 0.00722 -0.0168 0.1825 1.0000 8.000 1.1263 0.01467 0.00789 -0.0134 0.1753 1.0000 8.500 1.1677 0.01501 0.00836 -0.0109 0.1697 1.0000 9.000 1.2050 0.01563 0.00896 -0.0078 0.1631 1.0000 9.500 1.2426 0.01628 0.00970 -0.0049 0.1577 1.0000 10.000 1.2800 0.01688 0.01038 -0.0019 0.1529 1.0000