XFOIL Version 6.94 Calculated polar for: NACA 4413 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.090 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0309 0.02756 0.01834 -0.0878 0.9069 0.1926 -2.750 0.0093 0.02684 0.01761 -0.0901 0.8987 0.2134 -2.250 0.0821 0.02564 0.01659 -0.0933 0.8813 0.2648 -2.000 0.1310 0.02479 0.01589 -0.0969 0.8767 0.3062 -1.750 0.1554 0.02442 0.01566 -0.0963 0.8639 0.3429 -1.500 0.2017 0.02338 0.01498 -0.0993 0.8591 0.4102 -1.250 0.2275 0.02279 0.01489 -0.0988 0.8480 0.4997 -1.000 0.2550 0.02137 0.01468 -0.0964 0.8419 0.7814 -0.750 0.3303 0.02059 0.01381 -0.1038 0.8388 1.0000 -0.500 0.3522 0.02079 0.01379 -0.1028 0.8253 1.0000 -0.250 0.3966 0.02036 0.01315 -0.1052 0.8199 1.0000 0.000 0.4178 0.02059 0.01323 -0.1040 0.8065 1.0000 0.250 0.4599 0.02017 0.01264 -0.1058 0.8008 1.0000 0.500 0.4804 0.02048 0.01283 -0.1045 0.7876 1.0000 0.750 0.5206 0.02009 0.01229 -0.1059 0.7816 1.0000 1.000 0.5406 0.02045 0.01258 -0.1046 0.7686 1.0000 1.250 0.5791 0.02012 0.01211 -0.1057 0.7624 1.0000 1.500 0.5985 0.02056 0.01250 -0.1043 0.7496 1.0000 1.750 0.6358 0.02028 0.01210 -0.1053 0.7432 1.0000 2.000 0.6547 0.02080 0.01259 -0.1038 0.7308 1.0000 2.250 0.6907 0.02059 0.01226 -0.1046 0.7242 1.0000 2.500 0.7094 0.02116 0.01284 -0.1032 0.7122 1.0000 2.750 0.7441 0.02102 0.01260 -0.1038 0.7053 1.0000 3.000 0.7630 0.02164 0.01322 -0.1025 0.6940 1.0000 3.250 0.7963 0.02157 0.01309 -0.1029 0.6866 1.0000 3.500 0.8160 0.02219 0.01373 -0.1017 0.6760 1.0000 3.750 0.8473 0.02222 0.01370 -0.1018 0.6682 1.0000 4.000 0.8687 0.02279 0.01429 -0.1008 0.6583 1.0000 4.250 0.8977 0.02295 0.01444 -0.1007 0.6499 1.0000 4.500 0.9214 0.02342 0.01492 -0.1000 0.6407 1.0000 4.750 0.9476 0.02372 0.01523 -0.0995 0.6316 1.0000 5.000 0.9741 0.02406 0.01559 -0.0991 0.6230 1.0000