XFOIL Version 6.94 Calculated polar for: GOEN 611 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1338 0.06042 0.05458 0.0305 1.0002 0.8266 -2.750 -0.1935 0.06002 0.05447 0.0407 1.0002 0.8002 -2.500 -0.2405 0.05857 0.05331 0.0468 1.0002 0.7627 -2.250 -0.2708 0.05665 0.05157 0.0414 1.0002 0.6890 -2.000 -0.0489 0.04815 0.04070 -0.0411 1.0002 0.2913 -1.750 -0.0165 0.04641 0.03856 -0.0440 1.0002 0.2631 -1.500 0.0094 0.04536 0.03723 -0.0452 1.0002 0.2485 -1.250 0.0365 0.04495 0.03635 -0.0464 1.0002 0.2343 -1.000 0.0557 0.04426 0.03562 -0.0466 1.0002 0.2275 -0.750 0.0782 0.04433 0.03537 -0.0475 1.0002 0.2178 -0.500 0.0968 0.04499 0.03581 -0.0480 1.0002 0.2125 -0.250 0.1116 0.04571 0.03652 -0.0487 1.0002 0.2095 0.000 0.1774 0.04517 0.03560 -0.0572 0.9828 0.2083 0.250 0.2886 0.04271 0.03263 -0.0712 0.9427 0.2116 0.500 0.4176 0.03770 0.02758 -0.0849 0.8907 0.2270 0.750 0.5004 0.03328 0.02331 -0.0879 0.8277 0.2628 1.000 0.5307 0.03057 0.02206 -0.0842 0.7684 0.5152 1.250 0.5529 0.02954 0.02090 -0.0779 0.7086 0.9998 1.500 0.5763 0.02993 0.02041 -0.0733 0.6470 0.9998 1.750 0.5979 0.03063 0.02024 -0.0692 0.5800 0.9998 2.000 0.6198 0.03170 0.02047 -0.0658 0.5111 0.9998 2.250 0.6445 0.03316 0.02120 -0.0639 0.4549 0.9998 2.500 0.6720 0.03476 0.02228 -0.0630 0.4179 0.9998 2.750 0.7007 0.03638 0.02359 -0.0629 0.3918 0.9998 3.000 0.7288 0.03796 0.02491 -0.0628 0.3722 0.9998 3.250 0.7564 0.03956 0.02629 -0.0627 0.3563 0.9998 3.500 0.7843 0.04104 0.02740 -0.0625 0.3440 0.9998 3.750 0.8110 0.04325 0.02985 -0.0631 0.3342 0.9998 4.000 0.8384 0.04513 0.03160 -0.0632 0.3270 0.9998 4.250 0.8640 0.04762 0.03417 -0.0636 0.3216 0.9998 4.500 0.8868 0.05063 0.03749 -0.0642 0.3171 0.9998 4.750 0.9086 0.05353 0.04057 -0.0646 0.3125 0.9998 5.000 0.9318 0.05586 0.04286 -0.0646 0.3077 0.9998