XFOIL Version 6.94 Calculated polar for: manu4/33 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 -0.0494 0.08374 0.06291 -0.0190 0.9999 1.0001 -2.750 -0.0500 0.08176 0.06118 -0.0179 0.9999 1.0001 -2.500 -0.0516 0.07977 0.05945 -0.0167 0.9999 1.0001 -2.250 -0.0545 0.07772 0.05769 -0.0154 0.9999 1.0001 -2.000 -0.0589 0.07564 0.05586 -0.0138 0.9999 1.0001 -1.750 -0.0645 0.07352 0.05403 -0.0120 0.9999 1.0001 -1.500 -0.0698 0.07143 0.05217 -0.0104 0.9999 1.0001 -1.250 -0.0713 0.06949 0.05030 -0.0096 0.9999 1.0001 -1.000 -0.0648 0.06796 0.04862 -0.0103 0.9999 1.0001 -0.750 -0.0467 0.06699 0.04728 -0.0129 0.9999 1.0001 -0.500 -0.0164 0.06664 0.04633 -0.0175 0.9999 1.0001 -0.250 0.0234 0.06685 0.04580 -0.0234 0.9999 1.0001 0.000 0.0673 0.06752 0.04559 -0.0296 0.9999 1.0001 0.250 0.1097 0.06845 0.04575 -0.0349 0.9999 1.0001 0.500 0.1470 0.06955 0.04622 -0.0388 0.9999 1.0001 0.750 0.1780 0.07083 0.04709 -0.0417 0.9999 1.0001 1.000 0.2016 0.07246 0.04851 -0.0437 0.9999 1.0001 1.250 0.2147 0.07479 0.05077 -0.0452 0.9999 1.0001 1.500 0.2147 0.07814 0.05408 -0.0463 0.9999 1.0001 1.750 0.2092 0.08197 0.05770 -0.0472 0.9999 1.0001 2.000 0.2077 0.08558 0.06094 -0.0482 0.9999 1.0001 2.250 0.2104 0.08887 0.06381 -0.0494 0.9999 1.0001 2.500 0.2158 0.09195 0.06644 -0.0505 0.9999 1.0001 2.750 0.2227 0.09491 0.06896 -0.0516 0.9999 1.0001 3.000 0.2308 0.09777 0.07138 -0.0526 0.9999 1.0001 3.250 0.2397 0.10054 0.07372 -0.0536 0.9999 1.0001 3.500 0.2492 0.10327 0.07602 -0.0545 0.9999 1.0001 3.750 0.2592 0.10594 0.07829 -0.0554 0.9999 1.0001 4.000 0.2695 0.10858 0.08054 -0.0563 0.9999 1.0001 4.250 0.2800 0.11121 0.08279 -0.0571 0.9999 1.0001 4.500 0.2908 0.11380 0.08502 -0.0579 0.9999 1.0001 4.750 0.3017 0.11639 0.08727 -0.0587 0.9999 1.0001 5.000 0.3128 0.11896 0.08951 -0.0594 0.9999 1.0001