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 = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -4.500 -0.3473 0.02654 0.01772 -0.0263 0.9410 0.1231 -4.000 -0.2593 0.02420 0.01546 -0.0318 0.9185 0.1455 -3.500 -0.1654 0.02180 0.01320 -0.0377 0.8967 0.1765 -3.000 -0.0864 0.01932 0.01114 -0.0404 0.8717 0.2337 -2.500 -0.0510 0.01592 0.01006 -0.0342 0.8392 0.6295 -2.000 0.0178 0.01587 0.01034 -0.0300 0.7961 0.8309 -1.500 0.1198 0.01663 0.01056 -0.0321 0.7273 0.8974 -1.000 0.2468 0.01759 0.01043 -0.0414 0.6105 0.9435 -0.500 0.3840 0.01785 0.00964 -0.0562 0.5242 0.9879 0.000 0.4520 0.01770 0.00902 -0.0601 0.4868 1.0000 0.500 0.4881 0.01791 0.00893 -0.0577 0.4636 1.0000 1.000 0.5256 0.01826 0.00900 -0.0555 0.4453 1.0000 1.500 0.5642 0.01872 0.00926 -0.0532 0.4301 1.0000 2.000 0.6038 0.01930 0.00967 -0.0510 0.4165 1.0000 2.500 0.6451 0.02006 0.01018 -0.0489 0.4051 1.0000 3.000 0.6839 0.02086 0.01102 -0.0464 0.3939 1.0000 3.500 0.7241 0.02188 0.01196 -0.0440 0.3840 1.0000 4.000 0.7637 0.02290 0.01298 -0.0415 0.3746 1.0000 4.500 0.8020 0.02425 0.01438 -0.0389 0.3666 1.0000 5.000 0.8401 0.02547 0.01565 -0.0361 0.3577 1.0000 5.500 0.8767 0.02711 0.01736 -0.0333 0.3506 1.0000 6.000 0.9090 0.02876 0.01923 -0.0299 0.3429 1.0000 6.500 0.9518 0.03043 0.02077 -0.0280 0.3360 1.0000 7.000 0.9691 0.03288 0.02371 -0.0229 0.3289 1.0000 7.500 1.0051 0.03469 0.02558 -0.0203 0.3222 1.0000 8.000 1.0258 0.03758 0.02872 -0.0161 0.3157 1.0000 8.500 1.0284 0.04144 0.03300 -0.0105 0.3091 1.0000 9.000 1.0809 0.04279 0.03424 -0.0101 0.3030 1.0000 9.500 0.9930 0.05226 0.04441 0.0015 0.2970 1.0000