XFOIL Version 6.94 Calculated polar for: HQ 2.0/10 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 0.2751 0.00680 0.00214 -0.0599 0.7820 0.7744 0.500 0.3293 0.00677 0.00211 -0.0592 0.7630 0.7974 1.000 0.3836 0.00675 0.00209 -0.0586 0.7424 0.8175 1.500 0.4374 0.00673 0.00209 -0.0580 0.7220 0.8407 2.000 0.4901 0.00671 0.00210 -0.0570 0.6981 0.8685 2.500 0.5408 0.00666 0.00216 -0.0555 0.6716 0.9111 3.000 0.6148 0.00670 0.00223 -0.0593 0.6348 0.9979 3.500 0.6641 0.00714 0.00234 -0.0581 0.5562 1.0000 4.000 0.7047 0.00834 0.00272 -0.0556 0.3929 1.0000 4.500 0.7452 0.00982 0.00339 -0.0536 0.2427 1.0000 5.000 0.7945 0.01052 0.00388 -0.0527 0.1932 1.0000 5.500 0.8384 0.01172 0.00459 -0.0511 0.1041 1.0000 6.000 0.8824 0.01290 0.00546 -0.0495 0.0493 1.0000 6.500 0.9273 0.01396 0.00637 -0.0479 0.0204 1.0000 7.000 0.9702 0.01519 0.00755 -0.0460 0.0035 1.0000 7.500 1.0147 0.01620 0.00870 -0.0442 0.0031 1.0000 8.000 1.0564 0.01739 0.01009 -0.0419 0.0030 1.0000 8.500 1.0948 0.01879 0.01170 -0.0393 0.0029 1.0000 9.000 1.1286 0.02044 0.01358 -0.0360 0.0029 1.0000 9.500 1.1549 0.02233 0.01568 -0.0317 0.0030 1.0000 10.000 1.1699 0.02437 0.01794 -0.0258 0.0031 1.0000 10.500 1.1802 0.02683 0.02062 -0.0200 0.0032 1.0000 11.000 1.1854 0.02998 0.02405 -0.0146 0.0033 1.0000 11.500 1.1869 0.03389 0.02826 -0.0101 0.0035 1.0000 12.000 1.1833 0.03881 0.03356 -0.0064 0.0036 1.0000 12.500 1.1726 0.04499 0.04015 -0.0041 0.0038 1.0000 13.000 1.1559 0.05241 0.04798 -0.0036 0.0040 1.0000 13.500 1.1315 0.06182 0.05778 -0.0056 0.0041 1.0000 14.000 1.1039 0.07314 0.06946 -0.0105 0.0041 1.0000 14.500 1.0728 0.08713 0.08379 -0.0183 0.0042 1.0000 15.000 1.0432 0.10302 0.10000 -0.0282 0.0042 1.0000 15.500 1.0127 0.12068 0.11792 -0.0391 0.0042 1.0000 16.000 0.9839 0.13925 0.13671 -0.0504 0.0042 1.0000 16.500 0.9507 0.16090 0.15854 -0.0626 0.0042 1.0000