XFOIL Version 6.94 Calculated polar for: OAF095 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.5351 0.00973 0.00363 -0.1305 0.8108 0.4538 0.500 0.5925 0.00966 0.00368 -0.1303 0.7914 0.4907 1.000 0.6498 0.00961 0.00375 -0.1301 0.7731 0.5335 1.500 0.7069 0.00951 0.00384 -0.1298 0.7522 0.5921 2.000 0.7616 0.00915 0.00396 -0.1289 0.7254 0.7354 2.500 0.8082 0.00878 0.00387 -0.1259 0.6976 1.0000 3.000 0.8649 0.00900 0.00397 -0.1255 0.6637 1.0000 3.500 0.9211 0.00929 0.00416 -0.1251 0.6250 1.0000 4.000 0.9765 0.00967 0.00443 -0.1245 0.5771 1.0000 4.500 1.0299 0.01029 0.00480 -0.1237 0.5054 1.0000 5.500 1.1186 0.01423 0.00701 -0.1207 0.1789 1.0000 6.000 1.1616 0.01635 0.00856 -0.1191 0.0937 1.0000 6.500 1.2062 0.01805 0.01014 -0.1173 0.0675 1.0000 7.000 1.2493 0.01974 0.01183 -0.1152 0.0550 1.0000 7.500 1.2898 0.02159 0.01373 -0.1127 0.0470 1.0000 8.000 1.3262 0.02384 0.01604 -0.1096 0.0417 1.0000 8.500 1.3633 0.02581 0.01808 -0.1068 0.0373 1.0000 9.000 1.3976 0.02855 0.02098 -0.1035 0.0342 1.0000 9.500 1.4321 0.03071 0.02330 -0.1004 0.0313 1.0000 10.000 1.4658 0.03457 0.02731 -0.0975 0.0291 1.0000 10.500 1.4934 0.03758 0.03072 -0.0935 0.0274 1.0000 11.000 1.5117 0.04058 0.03401 -0.0887 0.0258 1.0000 11.500 1.5269 0.04427 0.03790 -0.0842 0.0247 1.0000 12.000 1.5299 0.05118 0.04525 -0.0795 0.0238 1.0000 12.500 1.5120 0.05717 0.05182 -0.0740 0.0235 1.0000 13.000 1.4844 0.06511 0.06032 -0.0705 0.0233 1.0000 13.500 1.4478 0.07506 0.07080 -0.0695 0.0232 1.0000 14.000 1.4037 0.08739 0.08361 -0.0720 0.0232 1.0000 14.500 1.3543 0.10249 0.09914 -0.0785 0.0234 1.0000 15.000 1.3041 0.12093 0.11795 -0.0897 0.0236 1.0000