XFOIL Version 6.96 Calculated polar for: MA409 (original) 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.500 -0.4291 0.09960 0.09627 -0.0290 1.0000 0.0471 -9.250 -0.4354 0.09726 0.09402 -0.0299 1.0000 0.0473 -9.000 -0.4427 0.09482 0.09166 -0.0303 1.0000 0.0474 -8.750 -0.4435 0.09161 0.08849 -0.0321 1.0000 0.0475 -8.500 -0.4432 0.08808 0.08501 -0.0342 1.0000 0.0476 -8.250 -0.4486 0.08261 0.07960 -0.0338 1.0000 0.0482 -8.000 -0.4476 0.07939 0.07642 -0.0316 1.0000 0.0488 -7.750 -0.4455 0.07656 0.07362 -0.0305 1.0000 0.0494 -7.500 -0.4436 0.07368 0.07076 -0.0303 1.0000 0.0500 -7.250 -0.4423 0.07072 0.06783 -0.0307 1.0000 0.0507 -7.000 -0.4117 0.06461 0.06165 -0.0398 0.9961 0.0517 -6.750 -0.3531 0.03710 0.03330 -0.0715 0.9892 0.0295 -6.500 -0.3137 0.02704 0.02217 -0.0798 0.9862 0.0302 -6.250 -0.2771 0.02138 0.01551 -0.0833 0.9847 0.0303 -6.000 -0.2413 0.01873 0.01231 -0.0853 0.9833 0.0320 -5.750 -0.2070 0.01701 0.01020 -0.0865 0.9811 0.0352 -5.500 -0.1744 0.01610 0.00916 -0.0877 0.9772 0.0424 -5.250 -0.1391 0.01550 0.00840 -0.0890 0.9740 0.0616 -5.000 -0.1032 0.01553 0.00830 -0.0909 0.9712 0.0843 -4.750 -0.0685 0.01555 0.00822 -0.0926 0.9679 0.0987 -4.500 -0.0374 0.01508 0.00767 -0.0934 0.9627 0.1066 -4.250 -0.0024 0.01488 0.00730 -0.0949 0.9593 0.1149 -4.000 0.0336 0.01413 0.00657 -0.0967 0.9572 0.1196 -3.750 0.0673 0.01368 0.00608 -0.0978 0.9535 0.1249 -3.500 0.0985 0.01322 0.00559 -0.0985 0.9486 0.1320 -3.250 0.1331 0.01287 0.00522 -0.0999 0.9456 0.1420 -3.000 0.1681 0.01238 0.00481 -0.1014 0.9434 0.1565 -2.750 0.1991 0.01185 0.00450 -0.1022 0.9395 0.1891 -2.500 0.2285 0.01126 0.00442 -0.1029 0.9351 0.3186 -2.250 0.2602 0.01088 0.00437 -0.1038 0.9317 0.4201 -2.000 0.2927 0.01043 0.00424 -0.1047 0.9289 0.5095 -1.750 0.3143 0.00972 0.00429 -0.1032 0.9232 0.7152 -1.500 0.3462 0.00923 0.00412 -0.1033 0.9194 1.0000 -1.250 0.3791 0.00917 0.00395 -0.1041 0.9155 1.0000 -1.000 0.4062 0.00918 0.00390 -0.1038 0.9080 1.0000 -0.750 0.4373 0.00912 0.00377 -0.1043 0.9027 1.0000 -0.500 0.4650 0.00913 0.00376 -0.1041 0.8957 1.0000 -0.250 0.4945 0.00911 0.00370 -0.1042 0.8897 1.0000 0.000 0.5222 0.00909 0.00367 -0.1040 0.8814 1.0000 0.250 0.5515 0.00900 0.00356 -0.1039 0.8719 1.0000 0.500 0.5797 0.00892 0.00349 -0.1035 0.8602 1.0000 0.750 0.6071 0.00889 0.00347 -0.1031 0.8478 1.0000 1.000 0.6351 0.00886 0.00343 -0.1028 0.8342 1.0000 1.250 0.6630 0.00883 0.00338 -0.1023 0.8162 1.0000 1.500 0.6912 0.00885 0.00334 -0.1018 0.7952 1.0000 1.750 0.7188 0.00901 0.00346 -0.1013 0.7751 1.0000 2.000 0.7451 0.00933 0.00365 -0.1005 0.7488 1.0000 2.250 0.7672 0.00967 0.00388 -0.0988 0.7053 1.0000 2.500 0.7862 0.00965 0.00367 -0.0965 0.5986 1.0000 2.750 0.8041 0.01056 0.00405 -0.0943 0.5481 1.0000 3.000 0.8261 0.01098 0.00438 -0.0931 0.4777 1.0000 3.250 0.8468 0.01155 0.00466 -0.0917 0.4012 1.0000 3.500 0.8635 0.01264 0.00515 -0.0898 0.2908 1.0000 3.750 0.8752 0.01477 0.00604 -0.0876 0.0842 1.0000 4.000 0.8949 0.01606 0.00699 -0.0862 0.0359 1.0000 4.250 0.9180 0.01684 0.00794 -0.0851 0.0318 1.0000 4.500 0.9397 0.01781 0.00904 -0.0838 0.0294 1.0000 4.750 0.9591 0.01910 0.01043 -0.0822 0.0280 1.0000 5.000 0.9760 0.02099 0.01238 -0.0802 0.0268 1.0000 5.250 0.9987 0.02197 0.01348 -0.0791 0.0252 1.0000 5.500 1.0206 0.02366 0.01526 -0.0778 0.0248 1.0000 5.750 1.0442 0.02571 0.01744 -0.0766 0.0248 1.0000 6.000 1.0687 0.02826 0.02020 -0.0756 0.0253 1.0000 6.250 1.0927 0.03304 0.02530 -0.0747 0.0269 1.0000 6.500 1.1187 0.03309 0.02566 -0.0730 0.0293 1.0000 13.500 0.9800 0.19058 0.18733 -0.1008 0.0369 1.0000 13.750 0.9794 0.19630 0.19301 -0.1072 0.0356 1.0000