XFOIL Version 6.96 Calculated polar for: HQ 3.5/8 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -7.500 -0.3610 0.09056 0.08736 -0.0300 1.0000 0.0275 -7.250 -0.3766 0.08957 0.08646 -0.0270 1.0000 0.0276 -7.000 -0.3922 0.08845 0.08542 -0.0246 1.0000 0.0278 -6.750 -0.4010 0.08660 0.08363 -0.0245 1.0000 0.0280 -6.500 -0.4056 0.08428 0.08135 -0.0259 1.0000 0.0282 -6.250 -0.4045 0.08129 0.07838 -0.0294 0.9999 0.0284 -6.000 -0.3667 0.07521 0.07216 -0.0418 0.9951 0.0287 -5.750 -0.3350 0.06978 0.06656 -0.0502 0.9902 0.0289 -5.500 -0.3240 0.06171 0.05856 -0.0540 0.9863 0.0303 -5.250 -0.3008 0.05812 0.05497 -0.0568 0.9831 0.0330 -5.000 -0.2696 0.05373 0.05043 -0.0629 0.9780 0.0373 -4.750 -0.2171 0.04813 0.04413 -0.0737 0.9726 0.0422 -4.500 -0.1916 0.04316 0.03928 -0.0769 0.9701 0.0444 -4.250 -0.1618 0.03998 0.03595 -0.0798 0.9655 0.0477 -3.750 -0.0869 0.03300 0.02827 -0.0872 0.9579 0.0585 -3.500 -0.0430 0.03077 0.02543 -0.0910 0.9557 0.0696 -3.250 -0.0188 0.02806 0.02275 -0.0917 0.9496 0.0736 -3.000 0.0179 0.02588 0.02027 -0.0941 0.9461 0.0869 -2.750 0.0673 0.02105 0.01429 -0.0936 0.9448 0.0299 -2.500 0.1074 0.01924 0.01222 -0.0955 0.9429 0.0280 -2.250 0.1369 0.01777 0.01054 -0.0955 0.9375 0.0280 -2.000 0.1717 0.01626 0.00897 -0.0968 0.9337 0.0332 -1.750 0.2112 0.01534 0.00797 -0.0989 0.9310 0.0395 -1.500 0.2539 0.01391 0.00664 -0.1018 0.9292 0.0919 -1.250 0.2725 0.01206 0.00688 -0.1005 0.9222 0.7005 -1.000 0.3018 0.01168 0.00675 -0.0996 0.9179 0.8160 -0.750 0.3295 0.01126 0.00646 -0.0985 0.9128 0.8931 -0.500 0.3772 0.01091 0.00602 -0.1022 0.9073 1.0000 -0.250 0.4193 0.01063 0.00558 -0.1048 0.9027 1.0000 0.000 0.4495 0.01053 0.00536 -0.1051 0.8932 1.0000 0.250 0.4846 0.01032 0.00503 -0.1062 0.8868 1.0000 0.500 0.5145 0.01020 0.00482 -0.1063 0.8775 1.0000 0.750 0.5430 0.01012 0.00466 -0.1061 0.8679 1.0000 1.000 0.5741 0.00998 0.00445 -0.1064 0.8606 1.0000 1.250 0.6013 0.00997 0.00440 -0.1060 0.8508 1.0000 1.500 0.6287 0.00996 0.00437 -0.1057 0.8411 1.0000 1.750 0.6571 0.00991 0.00427 -0.1055 0.8321 1.0000 2.000 0.6851 0.00986 0.00419 -0.1052 0.8219 1.0000 2.250 0.7116 0.00986 0.00418 -0.1047 0.8100 1.0000 2.500 0.7383 0.00985 0.00417 -0.1042 0.7975 1.0000 2.750 0.7650 0.00983 0.00419 -0.1036 0.7840 1.0000 3.000 0.7918 0.00983 0.00418 -0.1031 0.7695 1.0000 3.250 0.8183 0.00984 0.00418 -0.1025 0.7533 1.0000 3.500 0.8438 0.00989 0.00424 -0.1017 0.7334 1.0000 3.750 0.8697 0.00994 0.00427 -0.1010 0.7116 1.0000 4.000 0.8947 0.01004 0.00440 -0.1001 0.6847 1.0000 4.250 0.9189 0.01019 0.00450 -0.0991 0.6513 1.0000 4.500 0.9421 0.01044 0.00464 -0.0979 0.6094 1.0000 4.750 0.9637 0.01083 0.00483 -0.0965 0.5554 1.0000 5.000 0.9831 0.01142 0.00514 -0.0948 0.4913 1.0000 5.250 1.0016 0.01218 0.00564 -0.0931 0.4314 1.0000 5.500 1.0213 0.01291 0.00616 -0.0917 0.3851 1.0000 5.750 1.0418 0.01359 0.00670 -0.0906 0.3494 1.0000 6.000 1.0616 0.01428 0.00721 -0.0894 0.3090 1.0000 6.250 1.0820 0.01493 0.00776 -0.0883 0.2743 1.0000 6.500 1.1017 0.01562 0.00831 -0.0872 0.2327 1.0000 6.750 1.1207 0.01641 0.00887 -0.0860 0.1741 1.0000 7.000 1.1222 0.01952 0.01074 -0.0826 0.0271 1.0000 7.250 1.1384 0.02097 0.01239 -0.0805 0.0188 1.0000 7.500 1.1506 0.02271 0.01433 -0.0780 0.0162 1.0000 7.750 1.1612 0.02459 0.01637 -0.0753 0.0153 1.0000 8.000 1.1750 0.02626 0.01818 -0.0731 0.0149 1.0000 8.250 1.1900 0.02821 0.02028 -0.0711 0.0146 1.0000 8.500 1.2078 0.03052 0.02274 -0.0695 0.0145 1.0000 8.750 1.2279 0.03330 0.02574 -0.0683 0.0146 1.0000 9.000 1.2471 0.03664 0.02938 -0.0671 0.0148 1.0000 9.250 1.2619 0.04093 0.03405 -0.0655 0.0152 1.0000 9.500 1.2764 0.04479 0.03822 -0.0640 0.0159 1.0000 9.750 1.2897 0.04603 0.03971 -0.0617 0.0166 1.0000 10.000 1.2870 0.05081 0.04531 -0.0571 0.0195 1.0000 10.250 1.2757 0.05595 0.05091 -0.0531 0.0216 1.0000 10.500 1.2605 0.06015 0.05539 -0.0494 0.0227 1.0000 10.750 1.2423 0.06483 0.06029 -0.0464 0.0237 1.0000