XFOIL Version 6.94 Calculated polar for: HQ 2.5/9 B 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 0.3615 0.00820 0.00348 -0.0821 0.8475 0.8357 0.500 0.4131 0.00812 0.00342 -0.0807 0.8345 0.8635 1.000 0.4638 0.00796 0.00329 -0.0789 0.8197 0.8977 1.500 0.5190 0.00774 0.00318 -0.0782 0.8026 0.9505 2.000 0.5874 0.00770 0.00309 -0.0809 0.7824 1.0000 2.500 0.6433 0.00776 0.00310 -0.0807 0.7570 1.0000 3.000 0.6962 0.00782 0.00307 -0.0796 0.7147 1.0000 3.500 0.7449 0.00805 0.00303 -0.0776 0.6233 1.0000 4.000 0.7860 0.00912 0.00336 -0.0747 0.4809 1.0000 4.500 0.8283 0.01033 0.00399 -0.0725 0.3764 1.0000 5.000 0.8734 0.01133 0.00459 -0.0710 0.2904 1.0000 5.500 0.9143 0.01279 0.00534 -0.0690 0.1623 1.0000 6.000 0.9555 0.01436 0.00640 -0.0671 0.0800 1.0000 6.500 1.0010 0.01544 0.00735 -0.0657 0.0510 1.0000 7.000 1.0455 0.01657 0.00840 -0.0641 0.0296 1.0000 7.500 1.0889 0.01775 0.00956 -0.0624 0.0198 1.0000 8.000 1.1322 0.01884 0.01075 -0.0606 0.0168 1.0000 8.500 1.1741 0.01997 0.01202 -0.0586 0.0139 1.0000 9.000 1.2147 0.02110 0.01331 -0.0565 0.0110 1.0000 9.500 1.2531 0.02232 0.01474 -0.0540 0.0071 1.0000 10.000 1.2801 0.02427 0.01677 -0.0500 0.0031 1.0000 10.500 1.3030 0.02617 0.01893 -0.0454 0.0029 1.0000 11.000 1.3219 0.02838 0.02143 -0.0407 0.0028 1.0000 11.500 1.3364 0.03101 0.02438 -0.0361 0.0027 1.0000 12.000 1.3459 0.03416 0.02789 -0.0318 0.0027 1.0000 12.500 1.3487 0.03810 0.03220 -0.0280 0.0027 1.0000 13.000 1.3451 0.04300 0.03747 -0.0252 0.0026 1.0000 13.500 1.3339 0.04925 0.04411 -0.0237 0.0027 1.0000 14.000 1.3161 0.05721 0.05247 -0.0243 0.0027 1.0000 14.500 1.2905 0.06780 0.06346 -0.0280 0.0027 1.0000 15.000 1.2580 0.08200 0.07806 -0.0356 0.0027 1.0000 15.500 1.2197 0.09953 0.09596 -0.0461 0.0027 1.0000 16.000 1.1744 0.12013 0.11689 -0.0584 0.0028 1.0000 16.500 1.1221 0.14444 0.14148 -0.0727 0.0029 1.0000 17.000 1.0744 0.16920 0.16641 -0.0868 0.0031 1.0000 17.500 1.0493 0.18849 0.18580 -0.0976 0.0032 1.0000 18.000 1.0274 0.20870 0.20606 -0.1083 0.0035 1.0000