XFOIL Version 6.94 Calculated polar for: FX 74-130 WP1 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 0.3071 0.01861 0.01381 -0.0784 0.7852 0.9292 0.500 0.3575 0.01846 0.01355 -0.0772 0.7824 0.9397 1.000 0.4202 0.01842 0.01341 -0.0791 0.7801 0.9427 1.500 0.4256 0.01997 0.01501 -0.0729 0.7628 0.9502 2.000 0.4868 0.01997 0.01498 -0.0746 0.7590 0.9520 2.500 0.5589 0.01941 0.01439 -0.0778 0.7566 0.9532 3.000 0.6284 0.01905 0.01404 -0.0806 0.7542 0.9538 3.500 0.6971 0.01878 0.01385 -0.0834 0.7522 0.9547 4.500 0.7947 0.01876 0.01404 -0.0827 0.7304 0.9609 5.500 0.9701 0.01432 0.00987 -0.0916 0.7183 0.9602 6.000 1.0391 0.01215 0.00777 -0.0927 0.6913 0.9615 6.500 1.0812 0.01176 0.00752 -0.0902 0.6544 0.9655 7.000 1.1124 0.01189 0.00735 -0.0855 0.5510 0.9716 7.500 1.1067 0.01448 0.00919 -0.0770 0.4117 1.0000 8.000 1.0869 0.01926 0.01300 -0.0693 0.2435 1.0000 8.500 1.0834 0.02354 0.01639 -0.0640 0.1002 1.0000 9.000 1.0958 0.02705 0.01959 -0.0603 0.0485 1.0000 9.500 1.1160 0.03000 0.02262 -0.0577 0.0373 1.0000 10.000 1.1404 0.03273 0.02553 -0.0553 0.0317 1.0000 10.500 1.1556 0.03638 0.02931 -0.0520 0.0258 1.0000 11.000 1.1840 0.03919 0.03239 -0.0499 0.0222 1.0000 11.500 1.2097 0.04236 0.03568 -0.0477 0.0185 1.0000 12.500 1.2592 0.04909 0.04302 -0.0442 0.0132 1.0000 13.000 1.2724 0.05319 0.04738 -0.0424 0.0112 1.0000 13.500 1.2599 0.06250 0.05739 -0.0384 0.0093 1.0000 14.000 1.2453 0.07020 0.06563 -0.0362 0.0090 1.0000 14.500 1.2174 0.08007 0.07604 -0.0355 0.0089 1.0000 15.000 1.1818 0.09195 0.08841 -0.0373 0.0090 1.0000 15.500 1.1418 0.10601 0.10292 -0.0425 0.0090 1.0000 16.000 1.0997 0.12281 0.12011 -0.0517 0.0091 1.0000 16.500 1.0494 0.14494 0.14252 -0.0660 0.0095 1.0000