XFOIL Version 6.94 Calculated polar for: GOA 1 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 ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1808 0.02925 0.01603 -0.0320 1.0000 0.1206 -2.750 -0.1514 0.02746 0.01394 -0.0308 1.0000 0.1223 -2.500 -0.1238 0.02589 0.01213 -0.0294 1.0000 0.1273 -2.250 -0.0973 0.02440 0.01064 -0.0281 1.0000 0.1377 -2.000 -0.0707 0.02291 0.00930 -0.0271 1.0000 0.1647 -1.750 -0.0489 0.01848 0.00801 -0.0243 1.0000 1.0000 -1.500 -0.0228 0.01849 0.00733 -0.0229 1.0000 1.0000 -1.250 0.0024 0.01853 0.00701 -0.0217 1.0000 1.0000 -1.000 0.0274 0.01860 0.00685 -0.0206 1.0000 1.0000 -0.750 0.0522 0.01867 0.00674 -0.0196 1.0000 1.0000 -0.500 0.0771 0.01876 0.00671 -0.0185 1.0000 1.0000 -0.250 0.1024 0.01884 0.00680 -0.0175 1.0000 1.0000 0.000 0.1288 0.01892 0.00690 -0.0162 1.0000 1.0000 0.250 0.1581 0.01894 0.00754 -0.0151 1.0000 1.0000 0.500 0.2341 0.02293 0.00843 -0.0210 0.3129 1.0000 0.750 0.2595 0.02421 0.00929 -0.0200 0.2908 1.0000 1.000 0.2866 0.02541 0.01025 -0.0193 0.2698 1.0000 1.250 0.3152 0.02666 0.01133 -0.0189 0.2533 1.0000 1.500 0.3463 0.02796 0.01263 -0.0188 0.2463 1.0000 1.750 0.3788 0.02946 0.01416 -0.0189 0.2431 1.0000 2.000 0.4120 0.03090 0.01590 -0.0191 0.2422 1.0000 2.250 0.4442 0.03254 0.01790 -0.0193 0.2427 1.0000 2.500 0.4761 0.03414 0.02000 -0.0195 0.2456 1.0000 2.750 0.5083 0.03610 0.02247 -0.0202 0.2511 1.0000 3.000 0.5387 0.03848 0.02518 -0.0207 0.2571 1.0000 3.250 0.5710 0.04090 0.02822 -0.0220 0.2678 1.0000 3.500 0.6027 0.04375 0.03158 -0.0236 0.2814 1.0000 3.750 0.6337 0.04694 0.03528 -0.0258 0.2969 1.0000 4.000 0.6614 0.05053 0.03922 -0.0278 0.3078 1.0000 4.250 0.6926 0.05438 0.04374 -0.0330 0.3325 1.0000 4.500 0.7279 0.05986 0.04987 -0.0430 0.3915 1.0000 4.750 0.7470 0.06915 0.05983 -0.0739 0.5461 1.0000 5.000 0.6910 0.07447 0.06493 -0.0862 0.6512 1.0000