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 = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0075 0.03562 0.02059 -0.0269 1.0000 1.0000 -2.750 0.0119 0.03524 0.02020 -0.0252 1.0000 1.0000 -2.500 0.0145 0.03493 0.01988 -0.0232 1.0000 1.0000 -2.250 0.0155 0.03468 0.01963 -0.0209 1.0000 1.0000 -2.000 0.0147 0.03448 0.01943 -0.0184 1.0000 1.0000 -1.750 0.0124 0.03431 0.01925 -0.0156 1.0000 1.0000 -1.500 0.0089 0.03415 0.01909 -0.0127 1.0000 1.0000 -1.250 0.0045 0.03399 0.01891 -0.0096 1.0000 1.0000 -1.000 -0.0005 0.03383 0.01873 -0.0065 1.0000 1.0000 -0.750 -0.0060 0.03365 0.01854 -0.0034 1.0000 1.0000 -0.500 -0.0117 0.03346 0.01833 -0.0001 1.0000 1.0000 -0.250 -0.0176 0.03324 0.01810 0.0031 1.0000 1.0000 0.000 -0.0230 0.03303 0.01786 0.0063 1.0000 1.0000 0.250 -0.0271 0.03284 0.01763 0.0093 1.0000 1.0000 0.500 -0.0282 0.03272 0.01744 0.0118 1.0000 1.0000 0.750 -0.0242 0.03273 0.01737 0.0136 1.0000 1.0000 1.000 -0.0157 0.03288 0.01742 0.0146 1.0000 1.0000 1.250 -0.0042 0.03317 0.01763 0.0151 1.0000 1.0000 1.500 0.0091 0.03359 0.01799 0.0153 1.0000 1.0000 1.750 0.0232 0.03414 0.01850 0.0152 1.0000 1.0000 2.000 0.0372 0.03483 0.01918 0.0149 1.0000 1.0000 2.250 0.0505 0.03569 0.02005 0.0144 1.0000 1.0000 2.500 0.0622 0.03679 0.02118 0.0136 1.0000 1.0000 2.750 0.0715 0.03820 0.02265 0.0125 1.0000 1.0000 3.000 0.0779 0.03998 0.02449 0.0110 1.0000 1.0000 3.250 0.0830 0.04205 0.02659 0.0091 1.0000 1.0000 3.500 0.0887 0.04423 0.02880 0.0071 1.0000 1.0000 3.750 0.2115 0.05286 0.03772 -0.0166 0.8948 1.0000 4.000 0.2755 0.05733 0.04234 -0.0254 0.8319 1.0000 4.250 0.3049 0.06023 0.04529 -0.0283 0.7934 1.0000 4.500 0.3339 0.06317 0.04828 -0.0308 0.7609 1.0000 4.750 0.3657 0.06625 0.05140 -0.0334 0.7318 1.0000 5.000 0.3775 0.06881 0.05398 -0.0338 0.7112 1.0000