XFOIL Version 6.96 Calculated polar for: GOE 445 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.000 -0.5351 0.08875 0.08424 -0.0032 1.0000 0.0514 -8.750 -0.5399 0.08402 0.07955 -0.0052 1.0000 0.0523 -8.250 -0.6446 0.08440 0.07977 -0.0085 1.0000 0.0480 -6.750 -0.6385 0.05445 0.04891 -0.0165 1.0000 0.0240 -6.500 -0.6309 0.05025 0.04452 -0.0158 1.0000 0.0223 -6.250 -0.6207 0.04600 0.03995 -0.0147 1.0000 0.0205 -6.000 -0.6076 0.04143 0.03485 -0.0129 1.0000 0.0181 -5.750 -0.5870 0.03908 0.03189 -0.0108 1.0000 0.0164 -5.500 -0.5727 0.03517 0.02764 -0.0093 1.0000 0.0159 -5.250 -0.5554 0.03195 0.02398 -0.0076 1.0000 0.0157 -5.000 -0.5358 0.02909 0.02067 -0.0058 1.0000 0.0155 -4.750 -0.5144 0.02655 0.01769 -0.0042 1.0000 0.0155 -4.500 -0.4914 0.02428 0.01499 -0.0027 1.0000 0.0158 -4.250 -0.4685 0.02213 0.01256 -0.0013 1.0000 0.0173 -4.000 -0.4468 0.02030 0.01064 -0.0001 1.0000 0.0195 -3.750 -0.4254 0.01879 0.00903 0.0015 1.0000 0.0207 -3.500 -0.4050 0.01753 0.00761 0.0033 1.0000 0.0227 -3.250 -0.3846 0.01647 0.00640 0.0048 1.0000 0.0279 -3.000 -0.3631 0.01557 0.00528 0.0063 1.0000 0.0352 -2.750 -0.3468 0.01359 0.00424 0.0081 1.0000 0.2062 -2.500 -0.3379 0.01099 0.00445 0.0133 1.0000 0.7974 -2.250 -0.2873 0.01109 0.00436 0.0104 1.0000 0.9006 -2.000 -0.2455 0.01110 0.00402 0.0077 1.0000 0.9281 -1.750 -0.2005 0.01113 0.00378 0.0042 1.0000 0.9539 -1.500 -0.1449 0.01116 0.00356 -0.0016 1.0000 0.9787 -1.250 -0.0895 0.01107 0.00323 -0.0077 1.0000 0.9971 -1.000 -0.0662 0.01095 0.00303 -0.0074 1.0000 1.0000 -0.750 -0.0500 0.01086 0.00288 -0.0055 1.0000 1.0000 -0.500 -0.0336 0.01079 0.00277 -0.0037 1.0000 1.0000 -0.250 -0.0168 0.01076 0.00271 -0.0018 1.0000 1.0000 0.000 0.0000 0.01074 0.00269 0.0000 1.0000 1.0000 0.250 0.0168 0.01076 0.00271 0.0018 1.0000 1.0000 0.500 0.0336 0.01079 0.00277 0.0037 1.0000 1.0000 0.750 0.0500 0.01086 0.00288 0.0055 1.0000 1.0000 1.000 0.0662 0.01096 0.00303 0.0074 1.0000 1.0000 1.250 0.0894 0.01107 0.00323 0.0077 0.9971 1.0000 1.500 0.1450 0.01116 0.00356 0.0016 0.9787 1.0000 1.750 0.2003 0.01113 0.00378 -0.0042 0.9539 1.0000 2.000 0.2453 0.01110 0.00402 -0.0077 0.9282 1.0000 2.250 0.2872 0.01109 0.00436 -0.0104 0.9008 1.0000 2.500 0.3382 0.01099 0.00445 -0.0134 0.7974 1.0000 2.750 0.3472 0.01359 0.00425 -0.0082 0.2057 1.0000 3.000 0.3634 0.01559 0.00529 -0.0064 0.0350 1.0000 3.250 0.3850 0.01648 0.00641 -0.0049 0.0278 1.0000 3.500 0.4054 0.01756 0.00764 -0.0033 0.0225 1.0000 3.750 0.4258 0.01880 0.00904 -0.0016 0.0207 1.0000 4.000 0.4471 0.02030 0.01064 0.0000 0.0195 1.0000 4.250 0.4690 0.02209 0.01253 0.0013 0.0174 1.0000 4.500 0.4917 0.02429 0.01500 0.0026 0.0158 1.0000 4.750 0.5147 0.02653 0.01767 0.0041 0.0156 1.0000 5.000 0.5361 0.02907 0.02065 0.0058 0.0155 1.0000 5.250 0.5556 0.03194 0.02397 0.0075 0.0157 1.0000 5.500 0.5729 0.03517 0.02763 0.0092 0.0159 1.0000 5.750 0.5873 0.03897 0.03177 0.0107 0.0164 1.0000 6.000 0.6077 0.04130 0.03470 0.0128 0.0180 1.0000 6.250 0.6208 0.04596 0.03992 0.0147 0.0205 1.0000 6.500 0.6309 0.05023 0.04450 0.0158 0.0224 1.0000 6.750 0.6386 0.05435 0.04879 0.0165 0.0241 1.0000 9.250 0.6415 0.10235 0.09753 0.0024 0.0437 1.0000 9.500 0.6373 0.10687 0.10199 -0.0027 0.0431 1.0000 9.750 0.6333 0.11113 0.10617 -0.0075 0.0405 1.0000