XFOIL Version 6.94 Calculated polar for: GOA 3 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 -0.1531 0.04750 0.02053 -0.0235 0.9999 1.0001 -2.750 -0.1294 0.04731 0.01940 -0.0230 0.9999 1.0001 -2.500 -0.1058 0.04717 0.01842 -0.0224 0.9999 1.0001 -2.250 -0.0823 0.04707 0.01759 -0.0218 0.9999 1.0001 -2.000 -0.0590 0.04700 0.01690 -0.0211 0.9999 1.0001 -1.750 -0.0357 0.04696 0.01624 -0.0203 0.9999 1.0001 -1.500 -0.0124 0.04696 0.01579 -0.0196 0.9999 1.0001 -1.250 0.0107 0.04698 0.01545 -0.0188 0.9999 1.0001 -1.000 0.0339 0.04704 0.01524 -0.0180 0.9999 1.0001 -0.750 0.0571 0.04712 0.01514 -0.0172 0.9999 1.0001 -0.500 0.0802 0.04723 0.01515 -0.0163 0.9999 1.0001 -0.250 0.1035 0.04737 0.01520 -0.0155 0.9999 1.0001 0.000 0.1268 0.04753 0.01545 -0.0147 0.9999 1.0001 0.250 0.1503 0.04772 0.01583 -0.0139 0.9999 1.0001 0.500 0.1739 0.04793 0.01634 -0.0131 0.9999 1.0001 0.750 0.1980 0.04816 0.01701 -0.0123 0.9999 1.0001 1.000 0.2227 0.04841 0.01785 -0.0116 0.9999 1.0001 1.250 0.2483 0.04869 0.01885 -0.0111 0.9999 1.0001 1.500 0.2754 0.04900 0.02015 -0.0110 0.9999 1.0001 1.750 0.3039 0.04942 0.02183 -0.0116 0.9999 1.0001 2.000 0.3300 0.05024 0.02401 -0.0135 0.9999 1.0001 2.250 0.3444 0.05202 0.02675 -0.0162 0.9999 1.0001 2.500 0.3452 0.05491 0.02977 -0.0190 0.9999 1.0001 2.750 0.3432 0.05811 0.03271 -0.0215 0.9999 1.0001 3.000 0.3441 0.06121 0.03553 -0.0238 0.9999 1.0001 3.250 0.3476 0.06419 0.03830 -0.0261 0.9999 1.0001 3.500 0.3525 0.06711 0.04102 -0.0283 0.9999 1.0001 3.750 0.3586 0.06999 0.04374 -0.0303 0.9999 1.0001 4.000 0.3655 0.07285 0.04646 -0.0322 0.9999 1.0001 4.250 0.3730 0.07571 0.04918 -0.0340 0.9999 1.0001 4.500 0.3810 0.07859 0.05193 -0.0357 0.9999 1.0001 4.750 0.3894 0.08148 0.05471 -0.0374 0.9999 1.0001 5.000 0.3980 0.08438 0.05752 -0.0391 0.9999 1.0001