XFOIL Version 6.94 Calculated polar for: HQ 3.0/8 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 0.2919 0.00686 0.00206 -0.0693 0.7917 0.6965 1.500 0.4554 0.00697 0.00213 -0.0667 0.7018 0.7854 2.000 0.5098 0.00702 0.00223 -0.0660 0.6783 0.8114 2.500 0.5615 0.00717 0.00226 -0.0646 0.6169 0.8413 3.000 0.6128 0.00723 0.00241 -0.0632 0.5816 0.8826 3.500 0.6668 0.00773 0.00253 -0.0627 0.4703 1.0000 4.000 0.7212 0.00828 0.00289 -0.0626 0.4167 1.0000 4.500 0.7723 0.00920 0.00334 -0.0620 0.3260 1.0000 5.000 0.8238 0.01001 0.00389 -0.0615 0.2635 1.0000 5.500 0.8760 0.01072 0.00444 -0.0609 0.2213 1.0000 6.000 0.9264 0.01159 0.00506 -0.0602 0.1734 1.0000 6.500 0.9771 0.01240 0.00575 -0.0594 0.1417 1.0000 7.000 1.0280 0.01334 0.00651 -0.0589 0.1034 1.0000 7.500 1.0758 0.01447 0.00750 -0.0578 0.0724 1.0000 8.500 1.1627 0.01718 0.00997 -0.0544 0.0333 1.0000 9.000 1.2087 0.01873 0.01171 -0.0531 0.0327 1.0000 9.500 1.2497 0.01985 0.01279 -0.0512 0.0220 1.0000 10.500 1.3111 0.02482 0.01810 -0.0447 0.0231 1.0000 11.000 1.3403 0.02656 0.01998 -0.0414 0.0214 1.0000 11.500 1.3582 0.02954 0.02306 -0.0373 0.0207 1.0000 12.500 1.3952 0.03438 0.02833 -0.0310 0.0179 1.0000 13.000 1.4014 0.03912 0.03324 -0.0278 0.0175 1.0000 13.500 1.4061 0.04375 0.03818 -0.0253 0.0173 1.0000 14.000 1.4040 0.04930 0.04415 -0.0246 0.0163 1.0000 14.500 1.3970 0.05592 0.05109 -0.0245 0.0159 1.0000 15.000 1.3851 0.06373 0.05921 -0.0260 0.0156 1.0000 15.500 1.3681 0.07284 0.06861 -0.0289 0.0154 1.0000 16.000 1.3511 0.08274 0.07864 -0.0330 0.0145 1.0000 16.500 1.3235 0.09549 0.09181 -0.0395 0.0150 1.0000