XFOIL Version 6.94 Calculated polar for: bbl1 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.3262 0.05254 0.04037 -0.0624 0.8819 0.2385 -2.750 0.3636 0.05278 0.04102 -0.0651 0.8660 0.4039 -2.500 0.3915 0.05383 0.04219 -0.0660 0.8503 0.4483 -2.250 0.4500 0.05374 0.04246 -0.0711 0.8386 0.5186 -2.000 0.4609 0.05515 0.04411 -0.0693 0.8214 0.5523 -1.750 0.4836 0.05620 0.04545 -0.0690 0.8072 0.6005 -1.500 0.5208 0.05616 0.04583 -0.0696 0.7949 0.6653 -1.250 0.5253 0.05792 0.04780 -0.0667 0.7797 0.6993 -1.000 0.5647 0.05721 0.04747 -0.0664 0.7694 0.7583 -0.750 0.5588 0.05953 0.04998 -0.0625 0.7538 0.7816 -0.500 0.6000 0.05819 0.04895 -0.0618 0.7442 0.8491 -0.250 0.5827 0.06114 0.05214 -0.0571 0.7290 0.8995 0.000 0.6216 0.06186 0.05272 -0.0610 0.7148 1.0016 0.250 0.6655 0.06275 0.05334 -0.0655 0.6999 1.0016 0.500 0.6662 0.06608 0.05655 -0.0647 0.6853 1.0016 0.750 0.7300 0.06520 0.05543 -0.0692 0.6721 1.0016 1.000 0.7307 0.06823 0.05839 -0.0676 0.6560 1.0016 1.250 0.7324 0.07135 0.06147 -0.0662 0.6408 1.0016 1.500 0.7861 0.07037 0.06036 -0.0680 0.6249 1.0016 1.750 0.8642 0.06632 0.05622 -0.0705 0.6069 1.0016 2.000 0.8207 0.07330 0.06325 -0.0658 0.5883 1.0016 2.250 0.8382 0.07455 0.06447 -0.0643 0.5685 1.0016 2.500 0.8806 0.07311 0.06300 -0.0635 0.5485 1.0016 2.750 1.0075 0.06352 0.05334 -0.0675 0.5270 1.0016 3.000 1.0834 0.05941 0.04913 -0.0697 0.5049 1.0016 3.250 1.1322 0.05771 0.04738 -0.0701 0.4819 1.0016 3.500 1.1669 0.05716 0.04676 -0.0692 0.4587 1.0016 3.750 1.1968 0.05729 0.04682 -0.0682 0.4368 1.0016 4.000 1.2255 0.05762 0.04710 -0.0672 0.4158 1.0016 4.250 1.2559 0.05783 0.04718 -0.0664 0.3957 1.0016 4.500 1.2390 0.06153 0.05113 -0.0612 0.3811 1.0016 4.750 1.2251 0.06528 0.05501 -0.0567 0.3693 1.0016 5.000 1.1615 0.07305 0.06303 -0.0488 0.3629 1.0016