XFOIL Version 6.94 Calculated polar for: AH 79-100 C AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.8977 0.00815 0.00324 -0.2161 0.7913 0.6578 0.500 0.9340 0.00771 0.00336 -0.2113 0.7788 1.0000 1.000 0.9940 0.00791 0.00336 -0.2122 0.7669 1.0000 1.500 1.0413 0.00807 0.00348 -0.2104 0.7539 1.0000 2.000 1.0951 0.00829 0.00359 -0.2101 0.7411 1.0000 2.500 1.1465 0.00849 0.00371 -0.2091 0.7277 1.0000 3.000 1.1927 0.00870 0.00393 -0.2071 0.7126 1.0000 3.500 1.2407 0.00894 0.00412 -0.2055 0.6968 1.0000 4.000 1.2847 0.00919 0.00434 -0.2030 0.6780 1.0000 4.500 1.3274 0.00948 0.00460 -0.2002 0.6594 1.0000 5.000 1.3631 0.00978 0.00491 -0.1959 0.6387 1.0000 5.500 1.3931 0.01015 0.00529 -0.1905 0.6106 1.0000 6.000 1.4144 0.01078 0.00578 -0.1834 0.5643 1.0000 6.500 1.4333 0.01173 0.00651 -0.1761 0.5107 1.0000 7.000 1.4469 0.01306 0.00755 -0.1682 0.4470 1.0000 7.500 1.4596 0.01467 0.00887 -0.1606 0.3826 1.0000 8.000 1.4717 0.01654 0.01043 -0.1534 0.3170 1.0000 8.500 1.4845 0.01862 0.01221 -0.1468 0.2547 1.0000 9.000 1.4908 0.02131 0.01452 -0.1398 0.1810 1.0000 9.500 1.4759 0.02583 0.01826 -0.1308 0.0707 1.0000 10.000 1.4702 0.03025 0.02235 -0.1237 0.0195 1.0000 10.500 1.4862 0.03319 0.02544 -0.1195 0.0167 1.0000 11.000 1.5017 0.03632 0.02874 -0.1157 0.0153 1.0000 11.500 1.5082 0.04047 0.03305 -0.1116 0.0144 1.0000 12.000 1.5181 0.04453 0.03733 -0.1083 0.0138 1.0000 12.500 1.5240 0.04927 0.04226 -0.1053 0.0132 1.0000 13.000 1.5273 0.05461 0.04778 -0.1028 0.0127 1.0000 13.500 1.5288 0.06052 0.05385 -0.1008 0.0124 1.0000 14.000 1.5299 0.06683 0.06032 -0.0992 0.0121 1.0000 14.500 1.5369 0.07279 0.06641 -0.0979 0.0119 1.0000 15.000 1.5538 0.07767 0.07146 -0.0969 0.0118 1.0000 15.500 1.5701 0.08276 0.07679 -0.0962 0.0117 1.0000 16.000 1.5784 0.08890 0.08326 -0.0960 0.0115 1.0000 16.500 1.5866 0.09530 0.08999 -0.0962 0.0114 1.0000 17.000 1.5905 0.10244 0.09748 -0.0968 0.0114 1.0000 17.500 1.5863 0.11076 0.10618 -0.0984 0.0115 1.0000 18.000 1.5742 0.12030 0.11612 -0.1013 0.0115 1.0000 18.500 1.5569 0.13086 0.12706 -0.1055 0.0116 1.0000 19.000 1.5348 0.14250 0.13907 -0.1114 0.0116 1.0000 19.500 1.5135 0.15442 0.15130 -0.1180 0.0118 1.0000 21.500 1.0618 0.31711 0.31580 -0.1957 0.0182 1.0000 22.000 1.0634 0.32955 0.32824 -0.1996 0.0162 1.0000 22.500 1.0685 0.33880 0.33752 -0.2024 0.0149 1.0000 23.000 1.0694 0.35253 0.35125 -0.2065 0.0138 1.0000