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 = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -2.750 0.0347 0.03016 0.01871 -0.0263 1.0000 1.0000 -2.500 0.0363 0.02993 0.01849 -0.0242 1.0000 1.0000 -2.250 0.0353 0.02978 0.01835 -0.0217 1.0000 1.0000 -2.000 0.0320 0.02968 0.01827 -0.0189 1.0000 1.0000 -1.750 0.0271 0.02961 0.01818 -0.0160 1.0000 1.0000 -1.500 0.0215 0.02953 0.01810 -0.0129 1.0000 1.0000 -1.250 0.0152 0.02943 0.01799 -0.0098 1.0000 1.0000 -1.000 0.0085 0.02929 0.01785 -0.0067 1.0000 1.0000 -0.750 0.0013 0.02913 0.01768 -0.0035 1.0000 1.0000 -0.500 -0.0064 0.02892 0.01747 -0.0003 1.0000 1.0000 -0.250 -0.0146 0.02866 0.01720 0.0030 1.0000 1.0000 0.000 -0.0232 0.02837 0.01690 0.0064 1.0000 1.0000 0.250 -0.0318 0.02803 0.01655 0.0098 1.0000 1.0000 0.500 -0.0388 0.02771 0.01620 0.0129 1.0000 1.0000 0.750 -0.0397 0.02753 0.01597 0.0151 1.0000 1.0000 1.000 -0.0320 0.02759 0.01596 0.0160 1.0000 1.0000 1.250 -0.0190 0.02787 0.01616 0.0161 1.0000 1.0000 1.500 -0.0036 0.02832 0.01656 0.0157 1.0000 1.0000 1.750 0.0124 0.02895 0.01716 0.0150 1.0000 1.0000 2.000 0.0277 0.02978 0.01799 0.0141 1.0000 1.0000 2.250 0.1028 0.03293 0.02128 0.0016 0.9642 1.0000 2.500 0.2869 0.03640 0.02528 -0.0237 0.8315 1.0000 2.750 0.4116 0.03547 0.02470 -0.0321 0.7296 1.0000 3.000 0.4600 0.03468 0.02394 -0.0308 0.6710 1.0000 3.250 0.4947 0.03424 0.02340 -0.0285 0.6269 1.0000 3.500 0.5220 0.03420 0.02325 -0.0261 0.5897 1.0000 3.750 0.5477 0.03457 0.02350 -0.0242 0.5590 1.0000 4.000 0.5761 0.03475 0.02352 -0.0222 0.5305 1.0000 4.250 0.5956 0.03592 0.02464 -0.0209 0.5055 1.0000 4.500 0.6204 0.03703 0.02567 -0.0200 0.4851 1.0000 4.750 0.6453 0.03808 0.02661 -0.0188 0.4649 1.0000 5.000 0.6636 0.03982 0.02835 -0.0180 0.4463 1.0000