XFOIL Version 6.94 Calculated polar for: NACA 16-015 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 0.0000 0.01592 0.01164 0.0000 0.8983 0.8983 0.500 0.0595 0.01549 0.01121 -0.0014 0.8953 0.9056 1.000 0.1509 0.01526 0.01101 -0.0098 0.8947 0.9059 2.000 0.3448 0.01274 0.00855 -0.0269 0.8733 0.9060 2.500 0.4164 0.01221 0.00805 -0.0300 0.8494 0.9075 3.000 0.4751 0.01202 0.00790 -0.0306 0.8173 0.9110 3.500 0.5149 0.01184 0.00762 -0.0270 0.7592 0.9180 4.000 0.5279 0.01312 0.00790 -0.0174 0.5394 0.9245 4.500 0.5195 0.01496 0.00867 -0.0044 0.3073 0.9335 5.000 0.5276 0.01617 0.00922 0.0048 0.1612 0.9423 5.500 0.5575 0.01730 0.00993 0.0093 0.0783 0.9485 6.000 0.5964 0.01810 0.01066 0.0120 0.0612 0.9535 6.500 0.6407 0.01890 0.01149 0.0134 0.0537 0.9577 7.000 0.6774 0.01963 0.01231 0.0164 0.0492 0.9634 7.500 0.7218 0.02092 0.01362 0.0174 0.0439 0.9669 8.000 0.7606 0.02159 0.01441 0.0199 0.0404 0.9723 8.500 0.8105 0.02303 0.01591 0.0195 0.0358 0.9747 9.000 0.8478 0.02388 0.01688 0.0220 0.0321 0.9803 9.500 0.8997 0.02545 0.01859 0.0211 0.0281 0.9821 10.000 0.9440 0.02673 0.01993 0.0216 0.0250 0.9854 10.500 0.9837 0.02847 0.02196 0.0231 0.0223 0.9890 11.000 1.0250 0.02984 0.02344 0.0237 0.0203 0.9917 11.500 1.0578 0.03260 0.02657 0.0258 0.0186 0.9950 12.000 1.0892 0.03469 0.02898 0.0275 0.0171 0.9977 12.500 1.1122 0.03686 0.03128 0.0302 0.0161 1.0000 13.000 1.0957 0.04073 0.03556 0.0396 0.0155 1.0000 13.500 1.0728 0.04454 0.03980 0.0490 0.0151 1.0000 14.000 1.0400 0.04939 0.04507 0.0580 0.0148 1.0000 14.500 0.9985 0.05551 0.05160 0.0655 0.0145 1.0000 15.000 0.9472 0.06371 0.06019 0.0704 0.0145 1.0000 15.500 0.8844 0.07569 0.07256 0.0704 0.0146 1.0000 16.000 0.8122 0.09414 0.09137 0.0625 0.0150 1.0000