XFOIL Version 6.96 Calculated polar for: RAE 102 AIRFOIL 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 ------ -------- --------- --------- -------- -------- -------- 0.250 0.0313 0.00928 0.00400 -0.0002 0.8701 0.8989 0.500 0.0635 0.00933 0.00404 -0.0007 0.8556 0.9127 0.750 0.0988 0.00941 0.00410 -0.0018 0.8410 0.9244 1.000 0.1370 0.00951 0.00417 -0.0036 0.8263 0.9343 1.250 0.1748 0.00961 0.00425 -0.0054 0.8110 0.9446 1.500 0.2132 0.00971 0.00432 -0.0073 0.7941 0.9543 1.750 0.2561 0.00977 0.00437 -0.0102 0.7742 0.9608 2.000 0.2943 0.00984 0.00440 -0.0121 0.7535 0.9706 2.250 0.3365 0.00985 0.00439 -0.0149 0.7287 0.9780 2.500 0.3768 0.00987 0.00438 -0.0174 0.7033 0.9867 2.750 0.4183 0.00987 0.00436 -0.0202 0.6751 0.9952 3.000 0.4510 0.00984 0.00430 -0.0214 0.6454 1.0000 3.250 0.4686 0.00984 0.00426 -0.0195 0.6165 1.0000 3.500 0.4865 0.00989 0.00424 -0.0176 0.5819 1.0000 3.750 0.5042 0.01002 0.00428 -0.0156 0.5364 1.0000 4.000 0.5205 0.01030 0.00432 -0.0134 0.4561 1.0000 4.250 0.5295 0.01144 0.00456 -0.0103 0.2697 1.0000 4.500 0.5393 0.01297 0.00526 -0.0078 0.1276 1.0000 4.750 0.5548 0.01397 0.00603 -0.0058 0.0890 1.0000 5.000 0.5721 0.01480 0.00679 -0.0039 0.0725 1.0000 5.250 0.5898 0.01564 0.00762 -0.0021 0.0625 1.0000 5.500 0.6060 0.01686 0.00877 -0.0001 0.0559 1.0000 5.750 0.6266 0.01750 0.00949 0.0013 0.0500 1.0000 6.000 0.6452 0.01872 0.01065 0.0029 0.0458 1.0000 6.250 0.6657 0.02033 0.01236 0.0044 0.0433 1.0000 6.500 0.6877 0.02149 0.01368 0.0057 0.0408 1.0000 6.750 0.7089 0.02243 0.01470 0.0068 0.0373 1.0000 7.000 0.7300 0.02404 0.01639 0.0079 0.0356 1.0000 7.250 0.7500 0.02678 0.01933 0.0091 0.0345 1.0000 7.500 0.7675 0.03008 0.02298 0.0106 0.0341 1.0000 7.750 0.7827 0.03329 0.02657 0.0124 0.0341 1.0000 8.000 0.7938 0.03618 0.02989 0.0146 0.0335 1.0000