XFOIL Version 6.94 Calculated polar for: GOE 282 (DAIMLER XIII) 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 ------- -------- --------- --------- -------- ------- ------- 0.000 0.5860 0.01046 0.00328 -0.1192 0.7159 0.2196 0.500 0.6398 0.01056 0.00321 -0.1186 0.6755 0.2404 1.000 0.6936 0.01056 0.00318 -0.1181 0.6376 0.2715 1.500 0.7405 0.00909 0.00324 -0.1162 0.6048 1.0000 2.000 0.7942 0.00948 0.00334 -0.1157 0.5755 1.0000 2.500 0.8478 0.00990 0.00354 -0.1152 0.5497 1.0000 3.000 0.8966 0.01051 0.00376 -0.1137 0.4940 1.0000 3.500 0.9374 0.01166 0.00413 -0.1111 0.3422 1.0000 4.500 1.0213 0.01475 0.00579 -0.1070 0.1239 1.0000 5.000 1.0580 0.01689 0.00733 -0.1040 0.0091 1.0000 5.500 1.1067 0.01761 0.00816 -0.1026 0.0096 1.0000 6.000 1.1543 0.01840 0.00907 -0.1010 0.0108 1.0000 6.500 1.2003 0.01928 0.01011 -0.0991 0.0125 1.0000 7.000 1.2451 0.02018 0.01117 -0.0969 0.0148 1.0000 7.500 1.2869 0.02127 0.01246 -0.0941 0.0167 1.0000 8.000 1.3264 0.02241 0.01379 -0.0909 0.0187 1.0000 8.500 1.3610 0.02369 0.01530 -0.0869 0.0208 1.0000 9.000 1.3875 0.02534 0.01718 -0.0816 0.0231 1.0000 10.000 1.4121 0.03091 0.02319 -0.0690 0.0262 1.0000 10.500 1.4136 0.03499 0.02746 -0.0631 0.0264 1.0000 11.000 1.4173 0.03946 0.03202 -0.0581 0.0260 1.0000 11.500 1.4393 0.04361 0.03610 -0.0539 0.0250 1.0000 12.000 1.4577 0.04765 0.04044 -0.0504 0.0234 1.0000 12.500 1.5215 0.05324 0.04594 -0.0498 0.0189 1.0000 13.000 1.4914 0.05864 0.05193 -0.0443 0.0181 1.0000 13.500 1.4854 0.06582 0.05945 -0.0417 0.0158 1.0000 14.000 1.3140 0.06998 0.06442 -0.0263 0.0205 1.0000 14.500 1.3024 0.07773 0.07240 -0.0246 0.0187 1.0000 15.000 1.2588 0.08641 0.08150 -0.0246 0.0186 1.0000 15.500 1.2146 0.09390 0.08936 -0.0250 0.0187 1.0000 16.000 1.1758 0.10246 0.09827 -0.0276 0.0189 1.0000 16.500 1.1418 0.11127 0.10739 -0.0310 0.0190 1.0000 17.000 1.1126 0.12017 0.11656 -0.0349 0.0191 1.0000