XFOIL Version 6.94 Calculated polar for: nacb632 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4263 0.02734 0.02017 0.0320 1.0000 0.7929 -2.750 -0.2256 0.03098 0.02279 0.0181 1.0000 0.9082 -2.500 -0.1908 0.03022 0.02182 0.0151 1.0000 0.9244 -2.250 -0.1559 0.02950 0.02092 0.0119 1.0000 0.9393 -2.000 -0.1214 0.02883 0.02009 0.0086 1.0000 0.9535 -1.750 -0.0837 0.02822 0.01933 0.0045 1.0000 0.9677 -1.500 -0.0403 0.02763 0.01861 -0.0008 1.0000 0.9821 -1.250 0.0061 0.02707 0.01793 -0.0068 1.0000 0.9962 -1.000 0.0143 0.02686 0.01769 -0.0064 1.0000 1.0000 -0.750 0.0070 0.02671 0.01753 -0.0034 1.0000 1.0000 -0.500 -0.0007 0.02651 0.01733 -0.0004 1.0000 1.0000 -0.250 -0.0093 0.02624 0.01706 0.0027 1.0000 1.0000 0.000 -0.0188 0.02591 0.01672 0.0060 1.0000 1.0000 0.250 -0.0294 0.02550 0.01630 0.0094 1.0000 1.0000 0.500 -0.0402 0.02503 0.01582 0.0128 1.0000 1.0000 0.750 -0.0478 0.02459 0.01536 0.0158 1.0000 1.0000 1.000 -0.0222 0.02502 0.01574 0.0133 0.9927 1.0000 1.250 0.0681 0.02670 0.01740 0.0005 0.9579 1.0000 1.500 0.1758 0.02795 0.01873 -0.0136 0.9174 1.0000 1.750 0.3289 0.02773 0.01876 -0.0316 0.8614 1.0000 2.000 0.4488 0.02535 0.01662 -0.0403 0.7958 1.0000 2.250 0.4849 0.02367 0.01492 -0.0367 0.7327 1.0000 2.500 0.4991 0.02271 0.01377 -0.0310 0.6641 1.0000 2.750 0.5115 0.02229 0.01299 -0.0261 0.5998 1.0000 3.000 0.5249 0.02235 0.01265 -0.0225 0.5488 1.0000 3.250 0.5378 0.02278 0.01278 -0.0197 0.5072 1.0000 3.500 0.5533 0.02332 0.01302 -0.0173 0.4742 1.0000 3.750 0.5730 0.02405 0.01344 -0.0157 0.4480 1.0000 4.000 0.5933 0.02498 0.01424 -0.0146 0.4238 1.0000 4.250 0.6165 0.02597 0.01498 -0.0138 0.4022 1.0000 4.500 0.6412 0.02713 0.01592 -0.0133 0.3830 1.0000 4.750 0.6656 0.02845 0.01717 -0.0129 0.3661 1.0000 5.000 0.6895 0.02988 0.01852 -0.0126 0.3498 1.0000