XFOIL Version 6.94 Calculated polar for: NACA4412c 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.080 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1094 0.03197 0.02300 -0.0511 1.0001 0.1445 -2.750 -0.0840 0.03102 0.02216 -0.0514 1.0001 0.1530 -2.500 -0.0399 0.02988 0.02077 -0.0546 0.9954 0.1662 -2.250 0.0241 0.02813 0.01902 -0.0613 0.9836 0.1876 -2.000 0.0862 0.02643 0.01743 -0.0675 0.9700 0.2158 -1.750 0.1518 0.02442 0.01563 -0.0739 0.9529 0.2556 -1.500 0.2219 0.02197 0.01354 -0.0805 0.9298 0.3205 -1.000 0.2962 0.01682 0.01043 -0.0775 0.8049 0.9999 -0.750 0.3208 0.01724 0.00928 -0.0738 0.5475 0.9999 -0.500 0.3448 0.01840 0.00932 -0.0729 0.4435 0.9999 -0.250 0.3728 0.01909 0.00946 -0.0728 0.4147 0.9999 0.250 0.4309 0.02017 0.00983 -0.0731 0.3867 0.9999 0.500 0.4602 0.02080 0.01009 -0.0732 0.3784 0.9999 0.750 0.4902 0.02128 0.01040 -0.0735 0.3715 0.9999 1.000 0.5201 0.02186 0.01075 -0.0738 0.3654 0.9999 1.250 0.5499 0.02259 0.01119 -0.0741 0.3602 0.9999 1.500 0.5799 0.02320 0.01168 -0.0745 0.3555 0.9999 1.750 0.6096 0.02378 0.01218 -0.0748 0.3504 0.9999 2.000 0.6393 0.02444 0.01271 -0.0752 0.3455 0.9999 2.250 0.6690 0.02523 0.01332 -0.0756 0.3416 0.9999 2.500 0.6989 0.02624 0.01416 -0.0761 0.3384 0.9999 2.750 0.7284 0.02687 0.01484 -0.0764 0.3358 0.9999 3.000 0.7578 0.02761 0.01560 -0.0768 0.3331 0.9999 3.250 0.7871 0.02841 0.01641 -0.0771 0.3303 0.9999 3.500 0.8160 0.02924 0.01724 -0.0775 0.3272 0.9999 3.750 0.8447 0.03014 0.01811 -0.0778 0.3240 0.9999 4.000 0.8732 0.03125 0.01914 -0.0782 0.3210 0.9999 4.250 0.9012 0.03245 0.02037 -0.0785 0.3186 0.9999 4.500 0.9289 0.03341 0.02150 -0.0787 0.3167 0.9999 4.750 0.9562 0.03449 0.02274 -0.0789 0.3149 0.9999 5.000 0.9831 0.03568 0.02409 -0.0791 0.3130 0.9999