XFOIL Version 6.94 Calculated polar for: AH21 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 0.2391 0.00762 0.00404 -0.0628 0.9504 1.0000 0.500 0.3034 0.00716 0.00354 -0.0643 0.9374 1.0000 1.500 0.4683 0.00595 0.00236 -0.0744 0.9076 1.0000 2.000 0.5267 0.00565 0.00208 -0.0743 0.8631 1.0000 2.500 0.5814 0.00579 0.00188 -0.0733 0.7608 1.0000 3.000 0.6023 0.00707 0.00212 -0.0656 0.5326 1.0000 3.500 0.6286 0.00888 0.00273 -0.0601 0.2671 1.0000 4.000 0.6663 0.01047 0.00338 -0.0572 0.0913 1.0000 4.500 0.7132 0.01125 0.00398 -0.0557 0.0519 1.0000 5.000 0.7604 0.01203 0.00458 -0.0543 0.0258 1.0000 5.500 0.8082 0.01278 0.00531 -0.0529 0.0178 1.0000 6.000 0.8561 0.01353 0.00613 -0.0515 0.0147 1.0000 6.500 0.9037 0.01431 0.00702 -0.0501 0.0123 1.0000 7.000 0.9514 0.01503 0.00785 -0.0488 0.0091 1.0000 7.500 0.9979 0.01594 0.00893 -0.0471 0.0051 1.0000 8.000 1.0408 0.01735 0.01047 -0.0450 0.0028 1.0000 8.500 1.0828 0.01896 0.01237 -0.0426 0.0025 1.0000 9.000 1.1221 0.02096 0.01475 -0.0399 0.0024 1.0000 9.500 1.1565 0.02355 0.01780 -0.0365 0.0024 1.0000 10.000 1.1816 0.02710 0.02194 -0.0320 0.0024 1.0000 10.500 1.1903 0.03212 0.02769 -0.0256 0.0025 1.0000 11.000 1.1707 0.03823 0.03454 -0.0161 0.0025 1.0000 11.500 1.1293 0.04560 0.04254 -0.0070 0.0026 1.0000 12.000 1.0800 0.05481 0.05223 -0.0034 0.0026 1.0000 12.500 1.0365 0.06716 0.06488 -0.0087 0.0026 1.0000 13.000 0.9728 0.09637 0.09423 -0.0273 0.0027 1.0000 13.500 0.9072 0.12204 0.11983 -0.0412 0.0028 1.0000