XFOIL Version 6.94 Calculated polar for: USA 27 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.5001 0.01158 0.00365 -0.0916 0.6641 0.0516 1.000 0.5951 0.01155 0.00331 -0.0878 0.6019 0.0707 1.500 0.6352 0.01099 0.00350 -0.0848 0.5693 0.3710 3.000 0.8607 0.01115 0.00428 -0.0981 0.4626 1.0000 3.500 0.8994 0.01161 0.00457 -0.0947 0.4355 1.0000 4.000 0.9304 0.01219 0.00479 -0.0899 0.3862 1.0000 4.500 0.9661 0.01271 0.00511 -0.0860 0.3552 1.0000 5.000 1.0005 0.01327 0.00543 -0.0820 0.3158 1.0000 5.500 1.0330 0.01397 0.00587 -0.0777 0.2669 1.0000 6.000 1.0548 0.01536 0.00669 -0.0718 0.1776 1.0000 6.500 1.0782 0.01672 0.00773 -0.0663 0.1264 1.0000 7.000 1.0811 0.01933 0.00974 -0.0577 0.0069 1.0000 7.500 1.1154 0.02025 0.01077 -0.0543 0.0065 1.0000 8.000 1.1482 0.02133 0.01200 -0.0509 0.0066 1.0000 8.500 1.1781 0.02265 0.01350 -0.0473 0.0068 1.0000 9.000 1.2036 0.02432 0.01539 -0.0434 0.0071 1.0000 9.500 1.2227 0.02652 0.01785 -0.0391 0.0073 1.0000 10.000 1.2437 0.02873 0.02026 -0.0356 0.0078 1.0000 10.500 1.2538 0.03196 0.02376 -0.0316 0.0084 1.0000 11.000 1.2515 0.03656 0.02861 -0.0279 0.0089 1.0000 11.500 1.2380 0.04286 0.03512 -0.0252 0.0092 1.0000 12.000 1.2424 0.04806 0.04058 -0.0242 0.0099 1.0000 12.500 1.2314 0.05537 0.04812 -0.0237 0.0107 1.0000 13.000 1.2181 0.06292 0.05578 -0.0232 0.0113 1.0000 13.500 1.2253 0.06805 0.06108 -0.0221 0.0128 1.0000 14.000 1.0199 0.08240 0.07639 -0.0136 0.0105 1.0000 14.500 1.0189 0.08762 0.08157 -0.0120 0.0114 1.0000