XFOIL Version 6.96 Calculated polar for: USA 28 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -5.000 -0.4757 0.02723 0.01857 -0.0436 0.9555 0.1263 -4.500 -0.4163 0.02606 0.01733 -0.0448 0.9428 0.1545 -4.000 -0.3611 0.02503 0.01656 -0.0453 0.9293 0.2308 -3.500 -0.3010 0.02561 0.01717 -0.0465 0.9145 0.2817 -3.000 -0.2402 0.02623 0.01761 -0.0475 0.8996 0.3177 -2.500 -0.1826 0.02686 0.01821 -0.0478 0.8853 0.3496 -1.500 -0.0563 0.02626 0.01743 -0.0504 0.8588 0.4131 -1.000 0.0123 0.02560 0.01672 -0.0526 0.8479 0.4525 -0.500 0.0673 0.02523 0.01631 -0.0524 0.8337 0.4859 0.000 0.1286 0.02475 0.01585 -0.0534 0.8195 0.5167 1.000 0.3265 0.02222 0.01370 -0.0686 0.8001 0.5985 1.500 0.5766 0.01973 0.01259 -0.1075 0.7935 1.0000 2.000 0.6471 0.01946 0.01212 -0.1097 0.7733 1.0000 2.500 0.7229 0.01902 0.01152 -0.1126 0.7532 1.0000 3.000 0.7716 0.01908 0.01151 -0.1104 0.7260 1.0000 3.500 0.8264 0.01912 0.01147 -0.1094 0.6995 1.0000 4.000 0.8912 0.01910 0.01134 -0.1103 0.6755 1.0000 4.500 0.9331 0.01974 0.01200 -0.1072 0.6492 1.0000 5.000 0.9904 0.02023 0.01243 -0.1070 0.6281 1.0000 5.500 1.0242 0.02115 0.01350 -0.1026 0.6050 1.0000 6.000 1.0694 0.02171 0.01413 -0.1001 0.5832 1.0000 6.500 1.1158 0.02228 0.01481 -0.0979 0.5634 1.0000 7.000 1.1469 0.02310 0.01588 -0.0929 0.5427 1.0000 7.500 1.1836 0.02293 0.01577 -0.0881 0.5128 1.0000 8.000 1.2060 0.02239 0.01508 -0.0801 0.4667 1.0000 8.500 1.2253 0.02293 0.01547 -0.0725 0.4234 1.0000 9.000 1.2278 0.02378 0.01641 -0.0622 0.3817 1.0000 9.500 1.2075 0.02466 0.01739 -0.0482 0.3301 1.0000 10.000 1.1646 0.02766 0.01970 -0.0335 0.2035 1.0000