XFOIL Version 6.94 Calculated polar for: pp14 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0291 0.06601 0.04465 -0.0206 0.9996 1.0004 -2.750 -0.0294 0.06482 0.04373 -0.0191 0.9996 1.0004 -2.500 -0.0300 0.06364 0.04284 -0.0176 0.9996 1.0004 -2.250 -0.0307 0.06247 0.04195 -0.0161 0.9996 1.0004 -2.000 -0.0304 0.06132 0.04107 -0.0148 0.9996 1.0004 -1.750 -0.0254 0.06026 0.04023 -0.0147 0.9996 1.0004 -1.500 -0.0115 0.05938 0.03952 -0.0167 0.9996 1.0004 -1.250 0.0150 0.05883 0.03906 -0.0214 0.9996 1.0004 -1.000 0.0553 0.05870 0.03893 -0.0292 0.9996 1.0004 -0.750 0.1090 0.05906 0.03922 -0.0398 0.9996 1.0004 -0.500 0.1730 0.05992 0.03995 -0.0524 0.9996 1.0004 -0.250 0.2437 0.06124 0.04108 -0.0663 0.9996 1.0004 0.000 0.3168 0.06294 0.04255 -0.0803 0.9996 1.0004 0.250 0.3886 0.06493 0.04426 -0.0936 0.9996 1.0004 0.500 0.4560 0.06710 0.04609 -0.1055 0.9996 1.0004 0.750 0.5170 0.06934 0.04789 -0.1155 0.9996 1.0004 1.000 0.5703 0.07154 0.04959 -0.1233 0.9996 1.0004 1.250 0.6160 0.07362 0.05111 -0.1289 0.9996 1.0004 1.500 0.6543 0.07556 0.05248 -0.1325 0.9996 1.0004 1.750 0.6860 0.07734 0.05368 -0.1343 0.9996 1.0004 2.000 0.7125 0.07902 0.05491 -0.1350 0.9996 1.0004 2.250 0.7359 0.08066 0.05627 -0.1351 0.9996 1.0004 2.500 0.7575 0.08232 0.05783 -0.1350 0.9996 1.0004 2.750 0.7780 0.08404 0.05963 -0.1348 0.9996 1.0004 3.000 0.7977 0.08582 0.06158 -0.1346 0.9996 1.0004 3.250 0.8167 0.08769 0.06374 -0.1344 0.9996 1.0004 3.500 0.8352 0.08967 0.06612 -0.1342 0.9996 1.0004 3.750 0.8530 0.09180 0.06877 -0.1342 0.9996 1.0004 4.000 0.8698 0.09417 0.07181 -0.1343 0.9996 1.0004 4.250 0.8850 0.09693 0.07532 -0.1347 0.9996 1.0004 4.500 0.8968 0.10044 0.07965 -0.1359 0.9996 1.0004 4.750 0.9018 0.10533 0.08530 -0.1382 0.9996 1.0004 5.000 0.8978 0.11199 0.09243 -0.1421 0.9996 1.0004