XFOIL Version 6.94 Calculated polar for: NACA M15 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 ------- -------- --------- --------- -------- ------- ------- 1.500 0.5721 0.00998 0.00406 -0.0689 0.6146 1.0000 2.000 0.6231 0.01008 0.00405 -0.0681 0.5978 1.0000 2.500 0.6735 0.01019 0.00406 -0.0671 0.5801 1.0000 3.000 0.7235 0.01034 0.00408 -0.0660 0.5612 1.0000 3.500 0.7729 0.01047 0.00418 -0.0649 0.5414 1.0000 4.000 0.8183 0.01055 0.00403 -0.0628 0.4967 1.0000 4.500 0.8626 0.01094 0.00417 -0.0608 0.4513 1.0000 5.000 0.9078 0.01139 0.00451 -0.0589 0.4202 1.0000 5.500 0.9509 0.01196 0.00490 -0.0568 0.3825 1.0000 6.000 0.9941 0.01251 0.00535 -0.0547 0.3493 1.0000 6.500 1.0310 0.01345 0.00596 -0.0517 0.2806 1.0000 7.000 1.0560 0.01524 0.00714 -0.0471 0.1795 1.0000 7.500 1.0652 0.01804 0.00917 -0.0405 0.0551 1.0000 8.000 1.0889 0.01971 0.01068 -0.0359 0.0214 1.0000 8.500 1.1092 0.02126 0.01221 -0.0308 0.0057 1.0000 9.000 1.1343 0.02275 0.01384 -0.0270 0.0052 1.0000 9.500 1.1579 0.02460 0.01587 -0.0237 0.0050 1.0000 10.000 1.1789 0.02690 0.01835 -0.0209 0.0049 1.0000 10.500 1.1970 0.02970 0.02135 -0.0185 0.0049 1.0000 11.000 1.2115 0.03307 0.02493 -0.0165 0.0049 1.0000 11.500 1.2216 0.03709 0.02918 -0.0149 0.0049 1.0000 12.000 1.2273 0.04178 0.03409 -0.0137 0.0050 1.0000 12.500 1.2300 0.04697 0.03949 -0.0129 0.0050 1.0000 13.000 1.2294 0.05270 0.04544 -0.0125 0.0050 1.0000 13.500 1.2267 0.05903 0.05197 -0.0125 0.0051 1.0000 14.000 1.2226 0.06584 0.05899 -0.0131 0.0052 1.0000 14.500 1.2177 0.07312 0.06648 -0.0141 0.0052 1.0000 15.000 1.2125 0.08064 0.07420 -0.0155 0.0053 1.0000 15.500 1.2072 0.08836 0.08211 -0.0170 0.0054 1.0000 16.000 1.2025 0.09599 0.08993 -0.0186 0.0056 1.0000 16.500 1.1999 0.10334 0.09746 -0.0202 0.0057 1.0000 17.000 1.1986 0.11051 0.10482 -0.0219 0.0059 1.0000 17.500 1.1987 0.11749 0.11199 -0.0236 0.0060 1.0000 18.000 1.1994 0.12442 0.11912 -0.0255 0.0062 1.0000 18.500 1.1984 0.13173 0.12663 -0.0278 0.0064 1.0000 19.000 1.1971 0.13915 0.13426 -0.0304 0.0065 1.0000 19.500 1.1950 0.14727 0.14260 -0.0341 0.0067 1.0000 20.000 1.1822 0.15916 0.15484 -0.0416 0.0068 1.0000 20.500 1.1607 0.17374 0.16986 -0.0509 0.0071 1.0000 21.000 1.1262 0.19335 0.18994 -0.0637 0.0073 1.0000 21.500 1.0398 0.23894 0.23592 -0.0892 0.0085 1.0000