XFOIL Version 6.94 Calculated polar for: NACA M18 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 0.2391 0.01113 0.00306 -0.0273 0.5743 0.0498 0.500 0.2937 0.01078 0.00277 -0.0269 0.5608 0.0746 1.000 0.3444 0.01003 0.00272 -0.0263 0.5469 0.2953 1.500 0.4243 0.00885 0.00341 -0.0305 0.5320 0.9721 2.000 0.5698 0.00946 0.00379 -0.0495 0.5156 0.9973 2.500 0.6363 0.00954 0.00377 -0.0521 0.5026 1.0000 3.000 0.6887 0.00963 0.00378 -0.0517 0.4892 1.0000 3.500 0.7410 0.00976 0.00384 -0.0513 0.4772 1.0000 4.000 0.7925 0.00992 0.00390 -0.0508 0.4609 1.0000 4.500 0.8436 0.01008 0.00401 -0.0502 0.4427 1.0000 5.000 0.8943 0.01030 0.00419 -0.0495 0.4293 1.0000 5.500 0.9435 0.01052 0.00438 -0.0486 0.4031 1.0000 6.000 0.9916 0.01083 0.00462 -0.0475 0.3755 1.0000 6.500 1.0385 0.01124 0.00493 -0.0463 0.3390 1.0000 7.000 1.0782 0.01230 0.00562 -0.0442 0.2610 1.0000 7.500 1.0960 0.01506 0.00748 -0.0392 0.1165 1.0000 8.000 1.1101 0.01737 0.00931 -0.0332 0.0334 1.0000 8.500 1.1386 0.01837 0.01032 -0.0290 0.0240 1.0000 9.000 1.1554 0.01979 0.01169 -0.0232 0.0069 1.0000 9.500 1.1667 0.02098 0.01301 -0.0164 0.0059 1.0000 10.000 1.1823 0.02271 0.01485 -0.0119 0.0056 1.0000 10.500 1.2002 0.02501 0.01731 -0.0090 0.0053 1.0000 11.000 1.2160 0.02797 0.02044 -0.0069 0.0049 1.0000 11.500 1.2273 0.03166 0.02432 -0.0053 0.0045 1.0000 12.000 1.2361 0.03588 0.02873 -0.0042 0.0043 1.0000 12.500 1.2412 0.04064 0.03372 -0.0034 0.0042 1.0000 13.000 1.2423 0.04594 0.03922 -0.0028 0.0042 1.0000 13.500 1.2394 0.05184 0.04533 -0.0026 0.0042 1.0000 14.000 1.2343 0.05839 0.05208 -0.0028 0.0041 1.0000 14.500 1.2277 0.06543 0.05932 -0.0035 0.0041 1.0000 15.000 1.2200 0.07287 0.06695 -0.0045 0.0041 1.0000 15.500 1.2125 0.08048 0.07475 -0.0058 0.0041 1.0000 16.000 1.2053 0.08822 0.08267 -0.0073 0.0041 1.0000 16.500 1.1993 0.09590 0.09056 -0.0090 0.0041 1.0000 17.000 1.1940 0.10359 0.09843 -0.0108 0.0041 1.0000 17.500 1.1904 0.11114 0.10617 -0.0128 0.0042 1.0000 18.000 1.1878 0.11866 0.11388 -0.0151 0.0043 1.0000 18.500 1.1851 0.12632 0.12176 -0.0177 0.0044 1.0000 19.000 1.1814 0.13437 0.13006 -0.0207 0.0045 1.0000 19.500 1.1752 0.14310 0.13906 -0.0245 0.0047 1.0000 20.000 1.1634 0.15349 0.14975 -0.0298 0.0048 1.0000 20.500 1.1478 0.16531 0.16189 -0.0365 0.0050 1.0000 21.000 1.1284 0.17888 0.17576 -0.0449 0.0052 1.0000 21.500 1.1051 0.19475 0.19191 -0.0552 0.0053 1.0000 22.000 1.0709 0.21633 0.21378 -0.0687 0.0055 1.0000