XFOIL Version 6.94 Calculated polar for: HQ 2.5/10 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 0.3442 0.00690 0.00217 -0.0753 0.7851 0.7776 0.500 0.3987 0.00688 0.00213 -0.0748 0.7672 0.7977 1.000 0.4528 0.00686 0.00212 -0.0742 0.7476 0.8186 1.500 0.5060 0.00684 0.00211 -0.0734 0.7259 0.8426 2.000 0.5579 0.00679 0.00213 -0.0723 0.7032 0.8743 2.500 0.6131 0.00668 0.00217 -0.0718 0.6749 0.9433 3.000 0.6781 0.00686 0.00224 -0.0739 0.6396 1.0000 3.500 0.7312 0.00715 0.00243 -0.0735 0.5967 1.0000 4.000 0.7816 0.00761 0.00266 -0.0725 0.5324 1.0000 4.500 0.8203 0.00904 0.00321 -0.0698 0.3643 1.0000 5.000 0.8669 0.01000 0.00384 -0.0685 0.2849 1.0000 5.500 0.9154 0.01079 0.00446 -0.0675 0.2371 1.0000 6.000 0.9604 0.01191 0.00519 -0.0661 0.1612 1.0000 6.500 1.0055 0.01300 0.00603 -0.0647 0.1064 1.0000 7.000 1.0399 0.01517 0.00756 -0.0618 0.0141 1.0000 7.500 1.0830 0.01647 0.00903 -0.0597 0.0061 1.0000 8.000 1.1232 0.01799 0.01080 -0.0573 0.0044 1.0000 8.500 1.1554 0.02013 0.01324 -0.0537 0.0036 1.0000 9.000 1.1837 0.02233 0.01568 -0.0497 0.0033 1.0000 9.500 1.1993 0.02503 0.01866 -0.0439 0.0032 1.0000 10.000 1.2097 0.02830 0.02224 -0.0381 0.0031 1.0000 10.500 1.2172 0.03232 0.02665 -0.0329 0.0030 1.0000 11.000 1.2182 0.03743 0.03220 -0.0282 0.0031 1.0000 11.500 1.2111 0.04347 0.03869 -0.0245 0.0032 1.0000 12.000 1.1899 0.05149 0.04720 -0.0222 0.0034 1.0000 12.500 1.1618 0.06104 0.05719 -0.0227 0.0035 1.0000 13.000 1.1303 0.07237 0.06892 -0.0264 0.0035 1.0000