XFOIL Version 6.96 Calculated polar for: Smoothed ATR airfoil coordinates obtained using 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -5.000 -0.2696 0.01177 0.00527 -0.0405 0.8376 0.0627 -4.500 -0.2216 0.01123 0.00467 -0.0384 0.8015 0.0710 -4.000 -0.1737 0.01084 0.00412 -0.0363 0.7532 0.0801 -3.500 -0.1254 0.01058 0.00369 -0.0344 0.7013 0.0923 -3.000 -0.0778 0.01034 0.00331 -0.0324 0.6432 0.1129 -2.500 -0.0313 0.01016 0.00303 -0.0303 0.5810 0.1545 -2.000 0.0142 0.00988 0.00279 -0.0282 0.5120 0.2323 -1.500 0.0498 0.00913 0.00258 -0.0247 0.4281 0.4780 -0.500 0.1367 0.00874 0.00272 -0.0190 0.3664 0.7403 0.000 0.1851 0.00876 0.00290 -0.0168 0.3506 0.8192 0.500 0.2343 0.00904 0.00321 -0.0146 0.3373 0.8865 1.000 0.2850 0.00933 0.00348 -0.0126 0.3263 0.9240 1.500 0.3438 0.00977 0.00380 -0.0125 0.3165 0.9463 2.000 0.4224 0.01026 0.00417 -0.0166 0.3066 0.9617 2.500 0.4949 0.01063 0.00443 -0.0199 0.2978 0.9699 3.000 0.5742 0.01096 0.00464 -0.0248 0.2897 0.9746 3.500 0.6502 0.01136 0.00496 -0.0290 0.2823 0.9829 4.000 0.7314 0.01160 0.00514 -0.0344 0.2749 0.9889 4.500 0.8032 0.01204 0.00546 -0.0380 0.2681 0.9948 5.000 0.8769 0.01219 0.00564 -0.0420 0.2622 0.9997 5.500 0.9196 0.01246 0.00588 -0.0397 0.2571 1.0000 6.000 0.9582 0.01295 0.00632 -0.0367 0.2521 1.0000 6.500 0.9975 0.01320 0.00663 -0.0337 0.2479 1.0000 7.000 1.0353 0.01350 0.00696 -0.0304 0.2432 1.0000 7.500 1.0710 0.01401 0.00743 -0.0268 0.2385 1.0000 8.000 1.1073 0.01442 0.00790 -0.0233 0.2344 1.0000 8.500 1.1439 0.01478 0.00833 -0.0198 0.2301 1.0000 9.000 1.1810 0.01521 0.00876 -0.0166 0.2257 1.0000 9.500 1.2191 0.01588 0.00944 -0.0137 0.2209 1.0000 10.000 1.2569 0.01627 0.00996 -0.0107 0.2167 1.0000