XFOIL Version 6.96 Calculated polar for: HQ 3.0/15 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -18.000 -0.7071 0.12412 0.12187 -0.0425 1.0000 0.0094 -17.750 -0.7481 0.11195 0.10956 -0.0489 1.0000 0.0093 -17.500 -0.7833 0.10144 0.09893 -0.0545 1.0000 0.0093 -17.250 -0.8219 0.09069 0.08804 -0.0605 1.0000 0.0091 -17.000 -0.8592 0.08033 0.07753 -0.0664 1.0000 0.0091 -16.750 -0.8923 0.07076 0.06779 -0.0721 1.0000 0.0090 -16.500 -0.9154 0.06296 0.05985 -0.0769 1.0000 0.0090 -16.250 -0.9360 0.05620 0.05294 -0.0811 1.0000 0.0090 -16.000 -0.9473 0.05135 0.04798 -0.0839 1.0000 0.0091 -15.750 -0.9613 0.04648 0.04298 -0.0864 1.0000 0.0090 -15.500 -0.9706 0.04252 0.03891 -0.0882 1.0000 0.0091 -15.250 -0.9761 0.03925 0.03553 -0.0894 1.0000 0.0091 -15.000 -0.9785 0.03649 0.03268 -0.0901 1.0000 0.0093 -14.750 -0.9820 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0.8698 0.00778 0.00267 -0.0843 0.4481 0.7246 4.500 0.8937 0.00800 0.00283 -0.0836 0.4287 0.7314 4.750 0.9172 0.00825 0.00301 -0.0829 0.4090 0.7382 5.000 0.9401 0.00849 0.00320 -0.0820 0.3878 0.7452 5.250 0.9627 0.00878 0.00340 -0.0811 0.3652 0.7526 5.500 0.9843 0.00906 0.00362 -0.0801 0.3437 0.7601 5.750 1.0062 0.00935 0.00385 -0.0791 0.3238 0.7683 6.000 1.0273 0.00963 0.00409 -0.0780 0.3058 0.7763 6.500 1.0696 0.01017 0.00456 -0.0757 0.2782 0.7943 6.750 1.0899 0.01045 0.00482 -0.0744 0.2656 0.8044 7.000 1.1101 0.01070 0.00506 -0.0731 0.2544 0.8157 7.250 1.1296 0.01091 0.00529 -0.0717 0.2438 0.8283 7.500 1.1461 0.01115 0.00554 -0.0696 0.2325 0.8441 7.750 1.1604 0.01138 0.00582 -0.0672 0.2216 0.8684 8.000 1.1878 0.01160 0.00617 -0.0676 0.2060 1.0000 8.250 1.2033 0.01200 0.00650 -0.0656 0.1925 1.0000 8.500 1.2210 0.01234 0.00682 -0.0640 0.1838 1.0000 8.750 1.2371 0.01274 0.00719 -0.0622 0.1750 1.0000 9.000 1.2522 0.01320 0.00759 -0.0603 0.1649 1.0000 9.250 1.2685 0.01362 0.00799 -0.0587 0.1561 1.0000 9.500 1.2832 0.01413 0.00845 -0.0568 0.1480 1.0000 9.750 1.2994 0.01458 0.00889 -0.0553 0.1402 1.0000 10.000 1.3140 0.01512 0.00941 -0.0535 0.1325 1.0000 10.250 1.3289 0.01566 0.00992 -0.0519 0.1252 1.0000 10.500 1.3429 0.01627 0.01051 -0.0502 0.1174 1.0000 10.750 1.3556 0.01696 0.01117 -0.0484 0.1092 1.0000 11.000 1.3698 0.01760 0.01180 -0.0469 0.1025 1.0000 11.250 1.3812 0.01842 0.01257 -0.0451 0.0946 1.0000 11.500 1.3943 0.01916 0.01331 -0.0436 0.0877 1.0000 11.750 1.4052 0.02006 0.01417 -0.0419 0.0800 1.0000 12.000 1.4152 0.02104 0.01512 -0.0402 0.0724 1.0000 12.250 1.4269 0.02195 0.01603 -0.0388 0.0670 1.0000 12.500 1.4357 0.02307 0.01712 -0.0372 0.0605 1.0000 12.750 1.4464 0.02410 0.01815 -0.0358 0.0555 1.0000 13.000 1.4538 0.02538 0.01940 -0.0342 0.0492 1.0000 13.250 1.4622 0.02662 0.02065 -0.0328 0.0445 1.0000 13.500 1.4706 0.02790 0.02193 -0.0316 0.0409 1.0000 13.750 1.4781 0.02929 0.02333 -0.0303 0.0368 1.0000 14.000 1.4852 0.03074 0.02479 -0.0291 0.0335 1.0000 14.250 1.4926 0.03222 0.02629 -0.0280 0.0304 1.0000 14.500 1.4966 0.03401 0.02808 -0.0268 0.0262 1.0000 14.750 1.4973 0.03614 0.03020 -0.0256 0.0205 1.0000 15.000 1.5005 0.03815 0.03222 -0.0247 0.0176 1.0000 15.250 1.5041 0.04018 0.03428 -0.0239 0.0151 1.0000 15.500 1.4948 0.04353 0.03761 -0.0229 0.0079 1.0000 15.750 1.4959 0.04597 0.04011 -0.0224 0.0066 1.0000 16.000 1.4934 0.04889 0.04309 -0.0219 0.0053 1.0000 16.250 1.4958 0.05137 0.04564 -0.0217 0.0051 1.0000 16.500 1.4941 0.05437 0.04872 -0.0216 0.0047 1.0000 16.750 1.4932 0.05738 0.05181 -0.0217 0.0045 1.0000 17.000 1.4922 0.06051 0.05503 -0.0219 0.0044 1.0000 17.250 1.4888 0.06399 0.05861 -0.0223 0.0042 1.0000 17.500 1.4858 0.06754 0.06225 -0.0228 0.0041 1.0000 17.750 1.4814 0.07139 0.06620 -0.0236 0.0040 1.0000 18.000 1.4774 0.07528 0.07018 -0.0245 0.0040 1.0000 18.250 1.4702 0.07971 0.07473 -0.0258 0.0039 1.0000 18.500 1.4655 0.08390 0.07901 -0.0271 0.0039 1.0000 18.750 1.4581 0.08862 0.08385 -0.0287 0.0039 1.0000 19.000 1.4461 0.09416 0.08952 -0.0307 0.0038 1.0000 19.250 1.4382 0.09915 0.09461 -0.0327 0.0038 1.0000