XFOIL Version 6.94 Calculated polar for: NACA M9 AIRFOIL 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 ------- -------- --------- --------- -------- ------- ------- 0.500 0.5482 0.01316 0.00497 -0.0614 0.4655 0.0906 1.000 0.6039 0.01313 0.00494 -0.0608 0.4545 0.1225 1.500 0.6602 0.01382 0.00585 -0.0601 0.4442 0.2154 2.000 0.7148 0.01496 0.00685 -0.0592 0.4339 0.2438 2.500 0.7679 0.01533 0.00744 -0.0584 0.4248 0.2925 3.000 0.8220 0.01530 0.00731 -0.0579 0.4139 0.3075 3.500 0.8765 0.01542 0.00743 -0.0574 0.4033 0.3171 4.000 0.9307 0.01537 0.00731 -0.0569 0.3914 0.3208 4.500 0.9841 0.01524 0.00726 -0.0564 0.3784 0.3278 5.000 1.0370 0.01534 0.00729 -0.0558 0.3621 0.3312 5.500 1.0893 0.01541 0.00734 -0.0552 0.3410 0.3292 6.000 1.1319 0.01626 0.00763 -0.0539 0.2436 0.3259 6.500 1.1766 0.01722 0.00840 -0.0526 0.2156 0.3228 7.000 1.2189 0.01829 0.00931 -0.0511 0.1903 0.3206 7.500 1.2605 0.01928 0.01018 -0.0494 0.1636 0.3192 8.000 1.2958 0.02052 0.01125 -0.0472 0.1286 0.3190 8.500 1.2862 0.02402 0.01426 -0.0400 0.0471 0.3187 9.500 1.2815 0.02998 0.02035 -0.0302 0.0042 0.3174 10.000 1.2939 0.03290 0.02346 -0.0281 0.0042 0.3169 10.500 1.3034 0.03637 0.02713 -0.0265 0.0044 0.3164 11.000 1.3096 0.04045 0.03143 -0.0253 0.0045 0.3161 11.500 1.3129 0.04517 0.03638 -0.0248 0.0047 0.3158 12.000 1.3147 0.05038 0.04182 -0.0248 0.0049 0.3156 12.500 1.3129 0.05633 0.04801 -0.0254 0.0053 0.3155 13.000 1.3077 0.06321 0.05514 -0.0266 0.0055 0.3154 13.500 1.2985 0.07120 0.06341 -0.0288 0.0058 0.3155 14.000 1.2907 0.07947 0.07193 -0.0314 0.0060 0.3155 14.500 1.2872 0.08747 0.08015 -0.0341 0.0064 0.3158 15.000 1.2769 0.09686 0.08980 -0.0376 0.0068 0.3159 15.500 1.2632 0.10724 0.10044 -0.0418 0.0071 0.3161 16.000 1.2476 0.11837 0.11182 -0.0466 0.0073 0.3164 16.500 1.2402 0.12826 0.12193 -0.0510 0.0077 0.3172 17.000 1.2350 0.13787 0.13176 -0.0555 0.0083 0.3180 17.500 1.2254 0.14847 0.14255 -0.0607 0.0087 0.3185 18.000 1.2170 0.15883 0.15305 -0.0661 0.0091 0.3191 18.500 1.2251 0.16579 0.16019 -0.0696 0.0099 0.3202 19.000 1.2325 0.17241 0.16683 -0.0731 0.0111 0.3213 19.500 1.2554 0.17578 0.17030 -0.0745 0.0130 0.3235 20.000 1.2972 0.17451 0.16906 -0.0727 0.0162 0.3275