XFOIL Version 6.94 Calculated polar for: HQ 2.5/12 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3376 0.00942 0.00403 -0.0734 0.7759 0.7704 0.500 0.3916 0.00934 0.00391 -0.0727 0.7578 0.7865 1.000 0.4452 0.00927 0.00385 -0.0719 0.7411 0.8024 1.500 0.4993 0.00922 0.00379 -0.0712 0.7249 0.8191 2.000 0.5527 0.00917 0.00372 -0.0704 0.7068 0.8378 2.500 0.6047 0.00910 0.00370 -0.0693 0.6863 0.8593 3.000 0.6549 0.00903 0.00370 -0.0678 0.6635 0.8865 4.000 0.7813 0.00903 0.00381 -0.0707 0.5979 1.0000 4.500 0.8311 0.00935 0.00399 -0.0698 0.5520 1.0000 5.000 0.8715 0.01011 0.00428 -0.0671 0.4576 1.0000 5.500 0.9087 0.01120 0.00488 -0.0641 0.3665 1.0000 6.000 0.9425 0.01253 0.00568 -0.0608 0.2776 1.0000 6.500 0.9837 0.01344 0.00642 -0.0586 0.2392 1.0000 7.000 1.0245 0.01431 0.00719 -0.0563 0.2022 1.0000 7.500 1.0619 0.01532 0.00803 -0.0536 0.1610 1.0000 8.000 1.0938 0.01654 0.00903 -0.0500 0.1177 1.0000 8.500 1.1175 0.01795 0.01022 -0.0451 0.0775 1.0000 9.000 1.1227 0.02055 0.01232 -0.0381 0.0136 1.0000 9.500 1.1400 0.02261 0.01445 -0.0330 0.0059 1.0000 10.000 1.1624 0.02445 0.01650 -0.0291 0.0052 1.0000 10.500 1.1819 0.02658 0.01887 -0.0252 0.0048 1.0000 11.000 1.1971 0.02916 0.02170 -0.0215 0.0046 1.0000 11.500 1.2080 0.03222 0.02500 -0.0181 0.0045 1.0000 12.000 1.2114 0.03612 0.02916 -0.0151 0.0044 1.0000 12.500 1.2095 0.04084 0.03413 -0.0127 0.0044 1.0000 13.000 1.2051 0.04628 0.03981 -0.0114 0.0044 1.0000 13.500 1.1964 0.05278 0.04656 -0.0112 0.0044 1.0000 14.000 1.1865 0.06014 0.05417 -0.0120 0.0044 1.0000 14.500 1.1774 0.06812 0.06240 -0.0139 0.0044 1.0000 15.000 1.1675 0.07688 0.07141 -0.0167 0.0045 1.0000 15.500 1.1587 0.08605 0.08085 -0.0202 0.0045 1.0000 16.000 1.1490 0.09592 0.09098 -0.0243 0.0046 1.0000 16.500 1.1376 0.10661 0.10198 -0.0293 0.0047 1.0000 17.000 1.1227 0.11859 0.11428 -0.0356 0.0048 1.0000 17.500 1.1058 0.13155 0.12756 -0.0431 0.0050 1.0000 18.000 1.0875 0.14534 0.14165 -0.0514 0.0051 1.0000 18.500 1.0669 0.16031 0.15690 -0.0607 0.0052 1.0000 19.000 1.0469 0.17585 0.17268 -0.0704 0.0054 1.0000 19.500 1.0259 0.19259 0.18961 -0.0806 0.0056 1.0000