XFOIL Version 6.94 Calculated polar for: HQ-2.0/9 9.0% smoothed 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.2701 0.00774 0.00300 -0.0600 0.8498 0.8580 0.500 0.3200 0.00758 0.00285 -0.0580 0.8285 0.8878 1.000 0.3726 0.00745 0.00276 -0.0567 0.8091 0.9288 2.000 0.5073 0.00740 0.00261 -0.0615 0.7611 1.0000 2.500 0.5612 0.00754 0.00267 -0.0610 0.7338 1.0000 3.000 0.6145 0.00769 0.00279 -0.0603 0.7001 1.0000 3.500 0.6666 0.00791 0.00294 -0.0593 0.6525 1.0000 4.000 0.7156 0.00834 0.00311 -0.0577 0.5682 1.0000 4.500 0.7600 0.00932 0.00354 -0.0556 0.4444 1.0000 5.000 0.8048 0.01046 0.00426 -0.0539 0.3337 1.0000 5.500 0.8506 0.01160 0.00507 -0.0526 0.2509 1.0000 6.000 0.8955 0.01288 0.00599 -0.0512 0.1609 1.0000 6.500 0.9359 0.01477 0.00741 -0.0492 0.0774 1.0000 7.000 0.9789 0.01631 0.00892 -0.0474 0.0528 1.0000 7.500 1.0216 0.01788 0.01064 -0.0454 0.0322 1.0000 8.000 1.0581 0.02020 0.01306 -0.0425 0.0189 1.0000 8.500 1.0861 0.02399 0.01709 -0.0385 0.0152 1.0000 9.000 1.1210 0.02746 0.02101 -0.0355 0.0138 1.0000 9.500 1.1496 0.03256 0.02669 -0.0320 0.0131 1.0000 10.000 1.1678 0.03758 0.03227 -0.0278 0.0122 1.0000 10.500 1.1761 0.04070 0.03570 -0.0230 0.0109 1.0000 11.000 1.1634 0.04598 0.04142 -0.0173 0.0105 1.0000 11.500 1.1376 0.05306 0.04899 -0.0138 0.0104 1.0000 12.000 1.0969 0.06350 0.05997 -0.0144 0.0108 1.0000 12.500 1.0476 0.07789 0.07483 -0.0212 0.0111 1.0000 13.000 0.9840 0.10140 0.09876 -0.0376 0.0119 1.0000 13.500 0.7449 0.11676 0.11454 -0.0483 0.0153 1.0000