XFOIL Version 6.94 Calculated polar for: pp13 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1730 0.05784 0.04484 0.0147 0.9998 0.9117 -2.750 -0.2094 0.05718 0.04443 0.0233 0.9998 0.8989 -2.500 -0.2455 0.05639 0.04389 0.0316 0.9998 0.8862 -2.250 -0.2695 0.05549 0.04316 0.0369 0.9998 0.8697 -2.000 -0.2829 0.05474 0.04252 0.0393 0.9998 0.8487 -1.750 -0.2761 0.05448 0.04222 0.0366 0.9998 0.8201 -1.500 -0.2408 0.05428 0.04189 0.0278 0.9998 0.7840 -1.250 -0.1928 0.05429 0.04168 0.0169 0.9998 0.7490 -1.000 -0.1358 0.05467 0.04174 0.0044 0.9998 0.7141 -0.750 -0.0801 0.05448 0.04133 -0.0063 0.9998 0.6837 -0.500 -0.0212 0.05490 0.04143 -0.0178 0.9998 0.6517 -0.250 0.0320 0.05494 0.04116 -0.0269 0.9998 0.6251 0.000 0.0827 0.05508 0.04103 -0.0350 0.9998 0.6054 0.250 0.1303 0.05492 0.04067 -0.0417 0.9998 0.5931 0.500 0.1786 0.05514 0.04064 -0.0485 0.9998 0.5839 0.750 0.2254 0.05518 0.04053 -0.0546 0.9998 0.5768 1.000 0.2734 0.05553 0.04070 -0.0609 0.9998 0.5698 1.250 0.3200 0.05582 0.04090 -0.0667 0.9998 0.5703 1.500 0.3668 0.05611 0.04120 -0.0723 0.9998 0.5786 1.750 0.4139 0.05654 0.04170 -0.0779 0.9998 0.5902 2.250 0.5152 0.05729 0.04322 -0.0901 0.9998 0.6455 2.500 0.5783 0.05704 0.04429 -0.0988 0.9998 0.7431 2.750 0.6216 0.05879 0.04597 -0.1025 0.9998 1.0002 3.000 0.6599 0.06054 0.04757 -0.1055 0.9998 1.0002 3.250 0.6935 0.06218 0.04903 -0.1071 0.9998 1.0002 3.500 0.7239 0.06372 0.05041 -0.1076 0.9998 1.0002 3.750 0.7510 0.06527 0.05199 -0.1074 0.9998 1.0002 4.000 0.7731 0.06747 0.05476 -0.1078 0.9998 1.0002 4.250 1.0920 0.05939 0.04555 -0.1237 0.4532 1.0002 4.500 1.1175 0.06302 0.04905 -0.1236 0.4357 1.0002 4.750 1.1420 0.06681 0.05278 -0.1235 0.4226 1.0002 5.000 1.1543 0.07154 0.05775 -0.1238 0.4148 1.0002