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 = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0030 0.04981 0.03908 -0.0203 0.9998 0.9991 -2.750 -0.0304 0.04917 0.03867 -0.0110 0.9998 0.9853 -2.500 -0.0693 0.04864 0.03835 -0.0010 0.9998 0.9698 -2.250 -0.1073 0.04793 0.03785 0.0085 0.9998 0.9554 -2.000 -0.1487 0.04707 0.03720 0.0184 0.9998 0.9422 -1.750 -0.1936 0.04600 0.03635 0.0290 0.9998 0.9322 -1.500 -0.2366 0.04485 0.03538 0.0390 0.9998 0.9213 -1.250 -0.2632 0.04407 0.03465 0.0449 0.9998 0.9023 -1.000 -0.2434 0.04470 0.03516 0.0392 0.9998 0.8589 -0.750 -0.1620 0.04759 0.03764 0.0180 0.9998 0.7780 -0.500 -0.0704 0.04995 0.03960 -0.0041 0.9998 0.6894 -0.250 0.0010 0.05118 0.04048 -0.0194 0.9998 0.6194 0.000 0.0611 0.05207 0.04105 -0.0308 0.9998 0.5669 0.250 0.1107 0.05214 0.04088 -0.0384 0.9998 0.5329 0.500 0.1610 0.05273 0.04117 -0.0461 0.9998 0.5003 0.750 0.2060 0.05280 0.04102 -0.0519 0.9998 0.4752 1.000 0.2508 0.05304 0.04102 -0.0575 0.9998 0.4527 1.250 0.2948 0.05348 0.04120 -0.0627 0.9998 0.4347 1.500 0.3373 0.05360 0.04124 -0.0673 0.9998 0.4273 1.750 0.3794 0.05400 0.04157 -0.0717 0.9998 0.4246 2.000 0.4215 0.05452 0.04205 -0.0761 0.9998 0.4240 2.250 0.4641 0.05515 0.04271 -0.0805 0.9998 0.4237 2.500 0.5062 0.05587 0.04356 -0.0847 0.9998 0.4278 2.750 0.5505 0.05662 0.04464 -0.0890 0.9998 0.4413 3.000 0.6003 0.05729 0.04593 -0.0939 0.9998 0.4662 3.250 0.6594 0.05780 0.04752 -0.0998 0.9998 0.5062 3.500 0.9945 0.04596 0.03172 -0.1161 0.3238 1.0002 3.750 1.0280 0.04816 0.03346 -0.1158 0.3070 1.0002 4.000 1.0660 0.05057 0.03548 -0.1161 0.2962 1.0002 4.250 1.1040 0.05286 0.03796 -0.1167 0.2904 1.0002 4.500 1.1404 0.05550 0.04085 -0.1173 0.2864 1.0002 4.750 1.1738 0.05845 0.04413 -0.1178 0.2845 1.0002 5.000 1.2039 0.06181 0.04791 -0.1182 0.2838 1.0002