XFOIL Version 6.94 Calculated polar for: EPPLER 417 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 0.3954 0.01312 0.00830 -0.1219 0.8869 0.6505 0.500 0.4728 0.01245 0.00765 -0.1263 0.8850 0.6556 1.000 0.5515 0.01167 0.00693 -0.1309 0.8826 0.6601 1.500 0.6179 0.01097 0.00631 -0.1327 0.8738 0.6643 2.000 0.6830 0.01039 0.00578 -0.1345 0.8635 0.6697 2.500 0.7506 0.00977 0.00528 -0.1365 0.8502 0.6735 3.000 0.8072 0.00941 0.00500 -0.1364 0.8269 0.6780 3.500 0.8688 0.00914 0.00468 -0.1373 0.7777 0.6831 4.000 0.9149 0.00969 0.00472 -0.1348 0.6599 0.6868 4.500 0.9364 0.01074 0.00530 -0.1278 0.5576 0.6920 5.000 0.9521 0.01224 0.00617 -0.1202 0.4275 0.6964 5.500 0.9756 0.01388 0.00716 -0.1147 0.2931 0.7007 6.000 1.0071 0.01555 0.00819 -0.1110 0.1764 0.7062 6.500 1.0391 0.01717 0.00935 -0.1074 0.0852 0.7101 7.000 1.0680 0.01924 0.01095 -0.1034 0.0144 0.7150 7.500 1.1056 0.02058 0.01237 -0.1006 0.0091 0.7200 8.000 1.1425 0.02187 0.01384 -0.0977 0.0081 0.7247 8.500 1.1741 0.02368 0.01579 -0.0942 0.0075 0.7304 9.000 1.2025 0.02557 0.01791 -0.0903 0.0073 0.7351 9.500 1.2273 0.02784 0.02037 -0.0861 0.0071 0.7407 10.000 1.2487 0.03045 0.02320 -0.0817 0.0070 0.7457 10.500 1.2690 0.03347 0.02645 -0.0774 0.0069 0.7515 11.000 1.2898 0.03681 0.03008 -0.0736 0.0070 0.7566 11.500 1.3098 0.04072 0.03432 -0.0700 0.0071 0.7627 12.000 1.3254 0.04535 0.03935 -0.0664 0.0072 0.7683 12.500 1.3323 0.05094 0.04537 -0.0627 0.0074 0.7740 13.000 1.3333 0.05804 0.05286 -0.0595 0.0076 0.7794 13.500 1.3209 0.06396 0.05925 -0.0564 0.0077 0.7859 14.000 1.2475 0.08056 0.07694 -0.0554 0.0089 0.7888 14.500 1.2111 0.09327 0.09008 -0.0594 0.0092 0.7934 15.000 1.1751 0.10796 0.10512 -0.0669 0.0093 0.7986 15.500 1.1415 0.12429 0.12175 -0.0773 0.0095 0.8035