XFOIL Version 6.96 Calculated polar for: NASA SC(2)-0503 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -6.000 -0.4730 0.02913 0.02712 -0.0297 1.0000 0.0048 -5.750 -0.4535 0.02418 0.02198 -0.0318 1.0000 0.0050 -5.500 -0.4317 0.02019 0.01781 -0.0335 1.0000 0.0052 -5.250 -0.4081 0.01685 0.01427 -0.0348 1.0000 0.0055 -5.000 -0.3823 0.01387 0.01105 -0.0360 1.0000 0.0059 -4.750 -0.3548 0.01126 0.00819 -0.0370 1.0000 0.0065 -4.500 -0.3261 0.00906 0.00572 -0.0378 1.0000 0.0072 -4.250 -0.2969 0.00735 0.00372 -0.0383 1.0000 0.0082 -4.000 -0.2689 0.00659 0.00278 -0.0380 1.0000 0.0099 -2.500 -0.0843 0.00657 0.00110 -0.0429 1.0000 0.0096 -2.250 -0.0566 0.00631 0.00086 -0.0429 1.0000 0.0288 -2.000 -0.0218 0.00464 0.00066 -0.0459 1.0000 0.4844 -1.750 0.0073 0.00413 0.00065 -0.0465 1.0000 0.6508 -1.500 0.0346 0.00394 0.00069 -0.0464 1.0000 0.7288 -1.250 0.0613 0.00381 0.00075 -0.0461 1.0000 0.7884 -1.000 0.0880 0.00376 0.00078 -0.0458 1.0000 0.8156 -0.750 0.1147 0.00373 0.00082 -0.0456 1.0000 0.8340 -0.500 0.1415 0.00371 0.00087 -0.0453 1.0000 0.8491 -0.250 0.1682 0.00369 0.00093 -0.0451 1.0000 0.8637 0.000 0.2189 0.00407 0.00081 -0.0498 0.7891 0.8735 0.250 0.2391 0.00571 0.00099 -0.0487 0.3684 0.8869 0.500 0.2629 0.00704 0.00121 -0.0485 0.0213 0.9012 1.000 0.3120 0.00740 0.00172 -0.0466 0.0057 0.9421 1.250 0.3351 0.00872 0.00334 -0.0451 0.0041 1.0000 1.500 0.3628 0.00987 0.00457 -0.0448 0.0043 1.0000 1.750 0.3918 0.01182 0.00674 -0.0438 0.0072 1.0000 2.000 0.4045 0.00567 0.00128 -0.0395 0.0102 1.0000 2.250 0.4306 0.00689 0.00278 -0.0387 0.0101 1.0000 2.500 0.4592 0.00709 0.00317 -0.0381 0.0089 1.0000 2.750 0.4857 0.00820 0.00450 -0.0372 0.0075 1.0000 3.000 0.5107 0.00997 0.00654 -0.0362 0.0066 1.0000 3.250 0.5347 0.01219 0.00902 -0.0352 0.0060 1.0000 3.500 0.5577 0.01477 0.01185 -0.0344 0.0056 1.0000 3.750 0.5796 0.01776 0.01505 -0.0338 0.0052 1.0000 4.000 0.6002 0.02112 0.01861 -0.0334 0.0049 1.0000