XFOIL Version 6.94 Calculated polar for: HQ 1.0/9 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 0.1380 0.00631 0.00184 -0.0304 0.8074 0.8075 0.500 0.1906 0.00626 0.00183 -0.0292 0.7847 0.8345 1.000 0.2435 0.00622 0.00181 -0.0283 0.7626 0.8606 1.500 0.2957 0.00619 0.00181 -0.0271 0.7385 0.8904 2.000 0.3491 0.00616 0.00184 -0.0260 0.7114 0.9299 2.500 0.4191 0.00623 0.00193 -0.0289 0.6727 0.9713 3.000 0.4963 0.00645 0.00201 -0.0336 0.6094 1.0000 3.500 0.5329 0.00753 0.00218 -0.0302 0.4127 1.0000 4.000 0.5777 0.00857 0.00265 -0.0287 0.2790 1.0000 4.500 0.6233 0.00974 0.00320 -0.0274 0.1571 1.0000 5.000 0.6710 0.01073 0.00386 -0.0264 0.0841 1.0000 5.500 0.7186 0.01175 0.00464 -0.0253 0.0355 1.0000 6.000 0.7654 0.01295 0.00570 -0.0240 0.0056 1.0000 6.500 0.8128 0.01408 0.00698 -0.0225 0.0036 1.0000 7.000 0.8586 0.01545 0.00859 -0.0208 0.0034 1.0000 7.500 0.9000 0.01734 0.01074 -0.0185 0.0033 1.0000 8.000 0.9358 0.01992 0.01360 -0.0156 0.0034 1.0000 8.500 0.9665 0.02352 0.01757 -0.0120 0.0035 1.0000 9.000 0.9965 0.02719 0.02170 -0.0086 0.0036 1.0000 9.500 0.9964 0.03575 0.03147 -0.0016 0.0047 1.0000 10.500 0.9423 0.04792 0.04456 0.0113 0.0052 1.0000 11.000 0.9071 0.05598 0.05292 0.0112 0.0054 1.0000 12.000 0.8371 0.08409 0.08166 -0.0074 0.0052 1.0000 12.500 0.7888 0.11075 0.10846 -0.0222 0.0052 1.0000 13.000 0.7699 0.12600 0.12366 -0.0296 0.0054 1.0000 15.000 0.7618 0.18046 0.17797 -0.0528 0.0119 1.0000 15.500 0.7727 0.18917 0.18667 -0.0572 0.0111 1.0000 16.000 0.7860 0.19673 0.19423 -0.0611 0.0105 1.0000 16.500 0.8034 0.20282 0.20032 -0.0636 0.0100 1.0000 17.000 0.8141 0.21183 0.20931 -0.0685 0.0099 1.0000 17.500 0.8258 0.22082 0.21828 -0.0737 0.0092 1.0000 18.000 0.8405 0.22814 0.22559 -0.0777 0.0085 1.0000 18.500 0.8558 0.23487 0.23232 -0.0816 0.0081 1.0000 19.000 0.8721 0.24093 0.23838 -0.0848 0.0077 1.0000 21.000 0.7855 0.27622 0.27438 -0.1144 0.0070 1.0000 21.500 0.7955 0.28474 0.28292 -0.1183 0.0065 1.0000 22.000 0.8056 0.29314 0.29134 -0.1222 0.0062 1.0000 22.500 0.8153 0.30158 0.29980 -0.1259 0.0059 1.0000 23.000 0.8247 0.30993 0.30817 -0.1296 0.0057 1.0000 23.500 0.8342 0.31825 0.31652 -0.1330 0.0055 1.0000 24.500 0.8512 0.33654 0.33485 -0.1405 0.0054 1.0000 25.000 0.8592 0.34584 0.34417 -0.1445 0.0054 1.0000 25.500 0.8672 0.35516 0.35351 -0.1483 0.0054 1.0000 26.000 0.8750 0.36441 0.36279 -0.1520 0.0053 1.0000 26.500 0.8825 0.37383 0.37223 -0.1557 0.0052 1.0000 27.000 0.8897 0.38326 0.38169 -0.1593 0.0051 1.0000 27.500 0.8965 0.39262 0.39107 -0.1628 0.0050 1.0000 28.000 0.9030 0.40213 0.40060 -0.1664 0.0048 1.0000 28.500 0.9090 0.41163 0.41013 -0.1698 0.0047 1.0000 29.000 0.9147 0.42121 0.41974 -0.1734 0.0045 1.0000 29.500 0.9200 0.43079 0.42934 -0.1768 0.0042 1.0000 30.000 0.9248 0.44032 0.43889 -0.1802 0.0039 1.0000 30.500 0.9292 0.44975 0.44836 -0.1836 0.0037 1.0000