XFOIL Version 6.94 Calculated polar for: AH 80-140 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 0.3320 0.01524 0.00686 -0.0472 0.5354 0.1399 0.500 0.3779 0.01582 0.00746 -0.0453 0.5308 0.1216 1.000 0.5244 0.01328 0.00753 -0.0634 0.5231 0.9506 1.500 0.4392 0.01408 0.00771 -0.0348 0.5228 0.7354 2.000 0.6295 0.01343 0.00742 -0.0625 0.5125 0.9500 2.500 0.5571 0.01545 0.00800 -0.0359 0.5129 0.4679 3.000 0.8100 0.01447 0.00831 -0.0761 0.5022 0.9946 3.500 0.8282 0.01281 0.00648 -0.0674 0.4854 0.9843 4.000 0.8554 0.01431 0.00814 -0.0630 0.4888 0.9715 4.500 0.9219 0.01296 0.00653 -0.0639 0.4739 0.9802 5.000 1.0005 0.01192 0.00534 -0.0680 0.4457 1.0000 5.500 1.0405 0.01416 0.00783 -0.0661 0.4651 1.0000 6.000 1.0793 0.01474 0.00871 -0.0631 0.4616 1.0000 6.500 1.1295 0.01333 0.00718 -0.0607 0.4428 1.0000 7.000 1.1705 0.01425 0.00820 -0.0585 0.4441 1.0000 7.500 1.2160 0.01426 0.00826 -0.0561 0.4329 1.0000 8.000 1.2559 0.01462 0.00876 -0.0535 0.4256 1.0000 8.500 1.2776 0.01408 0.00755 -0.0465 0.3441 1.0000 9.000 1.2743 0.01615 0.00907 -0.0373 0.2622 1.0000 9.500 1.3597 0.01512 0.00910 -0.0416 0.3415 1.0000 10.000 1.3921 0.01595 0.00994 -0.0380 0.3245 1.0000 10.500 1.3793 0.01807 0.01181 -0.0279 0.2703 1.0000 11.000 1.1451 0.03177 0.02264 0.0018 0.1563 0.0809 12.000 1.1696 0.04894 0.04092 -0.0186 0.0317 1.0000 12.500 1.3445 0.03741 0.03107 -0.0189 0.1657 1.0000 13.000 1.2851 0.04752 0.04082 -0.0182 0.1137 1.0000 14.500 1.3025 0.06170 0.05533 -0.0176 0.0977 1.0000