XFOIL Version 6.94 Calculated polar for: nabab 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3794 0.04844 0.03965 -0.0058 0.9999 0.1385 -2.750 -0.3540 0.04549 0.03607 -0.0045 0.9999 0.1253 -2.500 -0.3314 0.04261 0.03322 -0.0031 0.9999 0.1196 -2.250 -0.2293 0.03718 0.02692 -0.0151 0.9424 0.1089 -2.000 -0.1635 0.03407 0.02305 -0.0185 0.6021 0.1080 -1.750 -0.1356 0.03431 0.02136 -0.0166 0.3961 0.1090 -1.500 -0.1008 0.03355 0.01994 -0.0167 0.3762 0.1109 -1.250 -0.0666 0.03266 0.01881 -0.0169 0.3650 0.1141 -1.000 -0.0345 0.03199 0.01791 -0.0167 0.3583 0.1176 -0.750 -0.0029 0.03135 0.01715 -0.0162 0.3536 0.1215 -0.500 0.0289 0.03084 0.01659 -0.0158 0.3502 0.1283 -0.250 0.0617 0.03029 0.01623 -0.0158 0.3478 0.1496 0.000 0.1564 0.02849 0.01677 -0.0284 0.3382 1.0001 0.250 0.1809 0.02894 0.01679 -0.0272 0.3040 1.0001 0.500 0.2047 0.02950 0.01682 -0.0261 0.2707 1.0001 0.750 0.2294 0.02985 0.01684 -0.0252 0.2457 1.0001 1.000 0.2539 0.03012 0.01679 -0.0243 0.2237 1.0001 1.250 0.2787 0.03033 0.01669 -0.0235 0.2058 1.0001 1.500 0.3034 0.03051 0.01654 -0.0227 0.1921 1.0001 1.750 0.3290 0.03041 0.01638 -0.0219 0.1803 1.0001 2.000 0.3543 0.03033 0.01610 -0.0212 0.1677 1.0001 2.250 0.3798 0.03019 0.01573 -0.0205 0.1538 1.0001 2.500 0.4062 0.03007 0.01542 -0.0198 0.1422 1.0001 2.750 0.4331 0.03002 0.01526 -0.0191 0.1325 1.0001 3.000 0.4600 0.03007 0.01518 -0.0182 0.1218 1.0001 3.250 0.4869 0.03022 0.01538 -0.0173 0.1126 1.0001 3.500 0.5142 0.03045 0.01586 -0.0164 0.1025 1.0001 3.750 0.5407 0.03081 0.01617 -0.0155 0.0840 1.0001 4.000 0.5664 0.03144 0.01638 -0.0147 0.0813 1.0001 4.250 0.5929 0.03207 0.01698 -0.0139 0.0806 1.0001 4.500 0.6196 0.03275 0.01772 -0.0130 0.0804 1.0001 4.750 0.6463 0.03350 0.01860 -0.0121 0.0806 1.0001 5.000 0.6740 0.03435 0.01974 -0.0113 0.0810 1.0001