XFOIL Version 6.94 Calculated polar for: pp14 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0033 0.04789 0.03844 -0.0168 0.9996 1.0004 -2.750 -0.0040 0.04691 0.03761 -0.0150 0.9996 1.0004 -2.500 -0.0052 0.04594 0.03680 -0.0132 0.9996 1.0004 -2.250 -0.0112 0.04501 0.03603 -0.0103 0.9996 0.9982 -2.000 -0.0520 0.04436 0.03556 0.0001 0.9996 0.9838 -1.750 -0.0932 0.04349 0.03487 0.0104 0.9996 0.9716 -1.500 -0.1199 0.04292 0.03443 0.0165 0.9996 0.9482 -1.250 0.2060 0.05304 0.04379 -0.0743 0.9996 0.6473 -1.000 0.3191 0.05470 0.04517 -0.1006 0.9996 0.5333 -0.750 0.3940 0.05548 0.04574 -0.1151 0.9996 0.4752 -0.500 0.4514 0.05582 0.04594 -0.1244 0.9996 0.4371 -0.250 0.4993 0.05609 0.04610 -0.1309 0.9996 0.4096 0.000 0.5436 0.05648 0.04638 -0.1362 0.9996 0.3892 0.250 0.5820 0.05678 0.04666 -0.1398 0.9996 0.3777 0.500 0.6189 0.05734 0.04718 -0.1429 0.9996 0.3733 0.750 0.6524 0.05789 0.04780 -0.1450 0.9996 0.3734 1.000 0.6846 0.05849 0.04852 -0.1467 0.9996 0.3745 1.250 0.7162 0.05913 0.04934 -0.1479 0.9996 0.3761 1.500 0.7477 0.05972 0.05024 -0.1487 0.9996 0.3825 2.000 1.0326 0.04373 0.03154 -0.1577 0.2832 1.0004 2.250 1.0616 0.04573 0.03252 -0.1562 0.2696 1.0004 2.500 1.0967 0.04764 0.03386 -0.1559 0.2589 1.0004 2.750 1.1418 0.04974 0.03550 -0.1572 0.2514 1.0004 3.000 1.1904 0.05174 0.03749 -0.1590 0.2466 1.0004 3.250 1.2410 0.05414 0.04000 -0.1615 0.2426 1.0004 3.500 1.2895 0.05691 0.04291 -0.1639 0.2384 1.0004 3.750 1.3336 0.06002 0.04624 -0.1659 0.2356 1.0004 4.000 1.3683 0.06317 0.04985 -0.1664 0.2359 1.0004 4.250 1.3942 0.06647 0.05380 -0.1657 0.2392 1.0004 4.500 1.4163 0.07029 0.05819 -0.1649 0.2434 1.0004 4.750 1.4358 0.07455 0.06290 -0.1640 0.2479 1.0004 5.000 1.4560 0.07930 0.06793 -0.1634 0.2523 1.0004