XFOIL Version 6.96 Calculated polar for: FX 71-L-150/20 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.000 -0.9293 0.15597 0.15389 0.0276 1.0000 0.0148 -18.750 -1.0322 0.12650 0.12396 0.0091 1.0000 0.0143 -18.500 -1.0611 0.11599 0.11326 0.0025 1.0000 0.0143 -18.250 -1.0940 0.10538 0.10241 -0.0037 1.0000 0.0141 -18.000 -1.1208 0.09648 0.09329 -0.0087 1.0000 0.0139 -17.750 -1.1389 0.08953 0.08617 -0.0124 1.0000 0.0139 -17.500 -1.1509 0.08387 0.08036 -0.0152 1.0000 0.0139 -17.250 -1.1685 0.07755 0.07382 -0.0180 1.0000 0.0136 -17.000 -1.1750 0.07315 0.06931 -0.0198 1.0000 0.0136 -16.750 -1.1794 0.06919 0.06525 -0.0214 1.0000 0.0137 -16.500 -1.1837 0.06536 0.06128 -0.0225 1.0000 0.0135 -16.250 -1.1850 0.06204 0.05787 -0.0236 1.0000 0.0136 -16.000 -1.1859 0.05884 0.05457 -0.0244 1.0000 0.0136 -15.750 -1.1859 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3.500 0.3898 0.00883 0.00256 0.0005 0.1824 0.6543 3.750 0.4152 0.00931 0.00282 0.0007 0.1319 0.6573 4.000 0.4407 0.00976 0.00309 0.0010 0.0908 0.6602 4.500 0.4935 0.01044 0.00355 0.0013 0.0472 0.6657 4.750 0.5195 0.01078 0.00381 0.0015 0.0295 0.6692 5.000 0.5463 0.01100 0.00403 0.0017 0.0238 0.6734 5.250 0.5728 0.01125 0.00426 0.0019 0.0175 0.6771 5.500 0.5993 0.01152 0.00451 0.0021 0.0159 0.6804 5.750 0.6259 0.01176 0.00476 0.0023 0.0153 0.6835 6.000 0.6524 0.01202 0.00503 0.0025 0.0151 0.6862 6.250 0.6785 0.01230 0.00532 0.0028 0.0148 0.6888 6.500 0.7042 0.01256 0.00563 0.0031 0.0146 0.6927 6.750 0.7296 0.01285 0.00595 0.0034 0.0145 0.6961 7.000 0.7546 0.01318 0.00631 0.0038 0.0143 0.6996 7.250 0.7794 0.01353 0.00668 0.0042 0.0142 0.7031 7.500 0.8037 0.01392 0.00709 0.0047 0.0142 0.7065 7.750 0.8274 0.01431 0.00752 0.0052 0.0141 0.7102 8.000 0.8506 0.01469 0.00795 0.0058 0.0140 0.7148 8.250 0.8730 0.01513 0.00843 0.0065 0.0140 0.7190 8.500 0.8948 0.01560 0.00894 0.0072 0.0139 0.7230 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0.0135 1.0000 14.250 1.1675 0.04175 0.03692 0.0256 0.0135 1.0000 14.500 1.1735 0.04370 0.03892 0.0256 0.0135 1.0000 14.750 1.1783 0.04579 0.04109 0.0256 0.0135 1.0000 15.000 1.1815 0.04810 0.04349 0.0257 0.0136 1.0000 15.250 1.1848 0.05044 0.04590 0.0255 0.0135 1.0000 15.500 1.1865 0.05307 0.04861 0.0252 0.0135 1.0000 15.750 1.1875 0.05589 0.05154 0.0249 0.0137 1.0000 16.000 1.1880 0.05882 0.05456 0.0241 0.0136 1.0000 16.250 1.1875 0.06195 0.05778 0.0232 0.0135 1.0000 16.500 1.1847 0.06548 0.06143 0.0224 0.0137 1.0000 16.750 1.1808 0.06924 0.06531 0.0211 0.0137 1.0000 17.000 1.1782 0.07296 0.06908 0.0195 0.0135 1.0000 17.250 1.1711 0.07744 0.07370 0.0177 0.0135 1.0000 17.500 1.1551 0.08350 0.07998 0.0151 0.0138 1.0000 17.750 1.1442 0.08895 0.08555 0.0124 0.0138 1.0000 18.000 1.1379 0.09375 0.09043 0.0099 0.0137 1.0000 18.250 1.1061 0.10346 0.10042 0.0045 0.0139 1.0000 18.500 1.0685 0.11488 0.11212 -0.0022 0.0142 1.0000 18.750 1.0331 0.12667 0.12413 -0.0094 0.0143 1.0000 19.000 0.9691 0.14616 0.14395 -0.0217 0.0146 1.0000