XFOIL Version 6.94 Calculated polar for: NASA SC(2)-0414 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.2883 0.00951 0.00506 -0.0822 0.8309 0.6935 0.500 0.3487 0.00952 0.00498 -0.0827 0.8035 0.7000 1.000 0.4046 0.00961 0.00501 -0.0820 0.7661 0.7030 1.500 0.4614 0.00976 0.00497 -0.0816 0.7032 0.7071 2.000 0.5129 0.01056 0.00497 -0.0806 0.5231 0.7110 2.500 0.5585 0.01217 0.00558 -0.0790 0.3037 0.7139 3.000 0.6110 0.01323 0.00603 -0.0788 0.1844 0.7180 3.500 0.6657 0.01396 0.00645 -0.0787 0.1332 0.7214 4.000 0.7195 0.01462 0.00699 -0.0783 0.1105 0.7244 4.500 0.7757 0.01517 0.00746 -0.0784 0.0945 0.7290 5.000 0.8288 0.01583 0.00805 -0.0778 0.0823 0.7323 5.500 0.8819 0.01650 0.00873 -0.0772 0.0737 0.7361 6.500 0.9853 0.01813 0.01037 -0.0756 0.0605 0.7436 7.000 1.0375 0.01886 0.01117 -0.0749 0.0561 0.7479 7.500 1.0871 0.01978 0.01209 -0.0738 0.0518 0.7517 8.500 1.1826 0.02191 0.01440 -0.0710 0.0455 0.7602 9.000 1.2256 0.02316 0.01570 -0.0689 0.0415 0.7635 11.000 1.3222 0.02700 0.02055 -0.0496 0.0346 0.7802 11.500 1.3416 0.02977 0.02356 -0.0448 0.0339 0.7841 12.000 1.3580 0.03300 0.02706 -0.0404 0.0333 0.7893 12.500 1.3669 0.03689 0.03124 -0.0358 0.0329 0.7930 13.000 1.3693 0.04152 0.03617 -0.0315 0.0325 0.7976 13.500 1.3630 0.04739 0.04237 -0.0279 0.0322 0.8016 14.000 1.3478 0.05461 0.04994 -0.0254 0.0320 0.8055 14.500 1.3223 0.06390 0.05959 -0.0249 0.0318 0.8101 15.000 1.2800 0.07644 0.07257 -0.0272 0.0318 0.8127 15.500 1.2068 0.09631 0.09298 -0.0363 0.0320 0.8145