XFOIL Version 6.94 Calculated polar for: GOE 673 AIRFOIL 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 0.2592 0.00752 0.00229 -0.0647 0.7260 0.6822 0.500 0.3146 0.00752 0.00236 -0.0643 0.6995 0.7299 1.000 0.3694 0.00758 0.00239 -0.0638 0.6694 0.7680 1.500 0.4225 0.00763 0.00248 -0.0628 0.6382 0.8140 2.000 0.4724 0.00767 0.00267 -0.0610 0.6131 0.8749 2.500 0.5216 0.00781 0.00284 -0.0590 0.5876 0.9385 3.000 0.5912 0.00803 0.00298 -0.0616 0.5592 0.9845 3.500 0.6613 0.00831 0.00318 -0.0648 0.5278 1.0000 4.000 0.7098 0.00864 0.00337 -0.0635 0.4921 1.0000 4.500 0.7607 0.00902 0.00364 -0.0626 0.4599 1.0000 5.000 0.8115 0.00947 0.00397 -0.0618 0.4223 1.0000 5.500 0.8621 0.01000 0.00436 -0.0608 0.3842 1.0000 6.000 0.9112 0.01063 0.00484 -0.0598 0.3394 1.0000 6.500 0.9579 0.01147 0.00544 -0.0584 0.2786 1.0000 7.000 0.9932 0.01329 0.00656 -0.0554 0.1573 1.0000 7.500 1.0183 0.01592 0.00850 -0.0510 0.0497 1.0000 8.000 1.0599 0.01704 0.00965 -0.0489 0.0423 1.0000 8.500 1.0967 0.01837 0.01103 -0.0460 0.0385 1.0000 9.000 1.1242 0.01995 0.01266 -0.0417 0.0354 1.0000 9.500 1.1485 0.02166 0.01446 -0.0373 0.0322 1.0000 10.000 1.1673 0.02386 0.01677 -0.0327 0.0310 1.0000 10.500 1.1913 0.02590 0.01895 -0.0293 0.0291 1.0000 11.000 1.2088 0.02859 0.02171 -0.0258 0.0278 1.0000 11.500 1.2250 0.03164 0.02489 -0.0225 0.0268 1.0000 12.000 1.2453 0.03442 0.02784 -0.0202 0.0248 1.0000 12.500 1.2620 0.03773 0.03119 -0.0177 0.0238 1.0000 13.000 1.2777 0.04131 0.03499 -0.0155 0.0228 1.0000 13.500 1.2910 0.04507 0.03896 -0.0142 0.0214 1.0000 14.000 1.3039 0.04903 0.04303 -0.0129 0.0206 1.0000 14.500 1.3085 0.05405 0.04829 -0.0120 0.0193 1.0000 15.000 1.3113 0.05956 0.05408 -0.0119 0.0184 1.0000 15.500 1.3131 0.06541 0.06013 -0.0125 0.0177 1.0000 16.000 1.3165 0.07113 0.06594 -0.0129 0.0170 1.0000 16.500 1.3029 0.07964 0.07482 -0.0147 0.0164 1.0000 17.000 1.2840 0.08978 0.08536 -0.0187 0.0157 1.0000 17.500 1.2722 0.09917 0.09499 -0.0229 0.0149 1.0000 18.000 1.2504 0.11043 0.10657 -0.0277 0.0148 1.0000 18.500 1.2486 0.11860 0.11478 -0.0322 0.0140 1.0000 19.000 1.2216 0.13181 0.12832 -0.0391 0.0140 1.0000 19.500 1.1865 0.14779 0.14463 -0.0487 0.0136 1.0000