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 = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1884 0.07039 0.06337 -0.0279 0.9998 0.2439 -2.750 -0.1842 0.06660 0.05970 -0.0239 0.9998 0.2502 -2.250 -0.1404 0.06357 0.05664 -0.0279 0.9998 0.2873 -2.000 -0.1229 0.06174 0.05484 -0.0280 0.9998 0.3083 -1.750 -0.1088 0.05949 0.05266 -0.0269 0.9998 0.3305 -1.500 -0.0968 0.05714 0.05039 -0.0250 0.9998 0.3545 -1.250 -0.0813 0.05574 0.04904 -0.0242 0.9998 0.3945 -0.250 -0.0656 0.04589 0.03976 -0.0042 0.9998 0.6179 0.000 -0.0544 0.04389 0.03785 -0.0010 0.9998 0.6661 0.250 -0.0342 0.04203 0.03604 0.0001 0.9998 0.7108 0.500 -0.0058 0.04040 0.03445 -0.0008 0.9998 0.7476 0.750 0.0353 0.03967 0.03372 -0.0053 0.9998 0.7745 1.000 0.0901 0.03966 0.03369 -0.0136 0.9998 0.7826 1.250 0.1472 0.04082 0.03480 -0.0234 0.9998 0.7589 1.500 0.1999 0.04349 0.03729 -0.0344 0.9998 0.6445 1.750 0.6799 0.03362 0.02157 -0.0959 0.1861 0.2204 2.000 0.7189 0.03443 0.02203 -0.0983 0.1717 0.2177 2.250 0.7611 0.03519 0.02253 -0.1010 0.1641 0.2169 2.500 0.8076 0.03591 0.02305 -0.1042 0.1583 0.2197 2.750 0.8731 0.03685 0.02375 -0.1109 0.1498 0.2364 3.000 0.9711 0.03745 0.02447 -0.1225 0.1458 0.2779 3.250 1.0630 0.03848 0.02619 -0.1331 0.1456 1.0002 3.500 1.0983 0.04034 0.02780 -0.1331 0.1465 1.0002 3.750 1.1311 0.04265 0.03005 -0.1330 0.1479 1.0002 4.000 1.1612 0.04439 0.03196 -0.1322 0.1500 1.0002 4.250 1.1893 0.04596 0.03392 -0.1309 0.1532 1.0002 4.500 1.2158 0.04831 0.03668 -0.1297 0.1568 1.0002 4.750 1.2402 0.05106 0.03978 -0.1284 0.1593 1.0002 5.000 1.2635 0.05433 0.04338 -0.1272 0.1636 1.0002