XFOIL Version 6.94 Calculated polar for: UNIVERSITY OF ILLINOIS UI-1720 AIRFOIL 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 0.0038 0.08368 0.06872 -0.0219 1.0004 0.9996 -2.750 -0.0095 0.08403 0.06914 -0.0206 1.0004 0.9996 -2.500 -0.0222 0.08421 0.06938 -0.0195 1.0004 0.9996 -2.250 -0.0344 0.08423 0.06945 -0.0184 1.0004 0.9996 -2.000 -0.0462 0.08411 0.06938 -0.0173 1.0004 0.9996 -1.750 -0.0580 0.08385 0.06916 -0.0163 1.0004 0.9996 -1.500 -0.0696 0.08344 0.06881 -0.0153 1.0004 0.9996 -1.250 -0.0813 0.08291 0.06832 -0.0143 1.0004 0.9996 -1.000 -0.0932 0.08223 0.06769 -0.0134 1.0004 0.9996 -0.750 -0.1052 0.08142 0.06693 -0.0124 1.0004 0.9996 -0.500 -0.1173 0.08046 0.06602 -0.0114 1.0004 0.9996 -0.250 -0.1294 0.07937 0.06497 -0.0105 1.0004 0.9996 0.000 -0.1408 0.07816 0.06377 -0.0099 1.0004 0.9996 0.250 -0.1471 0.07696 0.06249 -0.0106 1.0004 0.9996 0.500 -0.1418 0.07603 0.06128 -0.0144 1.0004 0.9996 0.750 -0.1170 0.07579 0.06041 -0.0230 1.0004 0.9996 1.000 -0.0703 0.07671 0.06015 -0.0365 1.0004 0.9996 1.250 -0.0214 0.07852 0.06035 -0.0486 1.0004 0.9996 1.500 0.0142 0.08044 0.06081 -0.0557 1.0004 0.9996 1.750 0.0413 0.08230 0.06144 -0.0598 1.0004 0.9996 2.000 0.0641 0.08412 0.06223 -0.0627 1.0004 0.9996 2.250 0.0846 0.08595 0.06314 -0.0649 1.0004 0.9996 2.500 0.1037 0.08779 0.06414 -0.0666 1.0004 0.9996 2.750 0.1219 0.08965 0.06526 -0.0682 1.0004 0.9996 3.000 0.1394 0.09153 0.06646 -0.0695 1.0004 0.9996 3.250 0.1564 0.09343 0.06774 -0.0707 1.0004 0.9996 3.500 0.1730 0.09537 0.06908 -0.0719 1.0004 0.9996 3.750 0.1892 0.09733 0.07050 -0.0729 1.0004 0.9996 4.000 0.2051 0.09932 0.07199 -0.0739 1.0004 0.9996 4.250 0.2208 0.10134 0.07353 -0.0749 1.0004 0.9996 4.500 0.2362 0.10338 0.07513 -0.0759 1.0004 0.9996 4.750 0.2515 0.10545 0.07677 -0.0768 1.0004 0.9996 5.000 0.2666 0.10754 0.07848 -0.0777 1.0004 0.9996