XFOIL Version 6.94 Calculated polar for: HQ-2.0/9 9.0% smoothed 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 0.2764 0.00644 0.00197 -0.0608 0.8099 0.8060 0.500 0.3305 0.00641 0.00193 -0.0601 0.7914 0.8291 1.000 0.3833 0.00635 0.00191 -0.0591 0.7708 0.8549 1.500 0.4349 0.00628 0.00190 -0.0578 0.7477 0.8908 2.000 0.4939 0.00618 0.00190 -0.0580 0.7227 0.9569 2.500 0.5598 0.00629 0.00196 -0.0603 0.6925 1.0000 3.000 0.6135 0.00651 0.00207 -0.0598 0.6533 1.0000 3.500 0.6643 0.00692 0.00224 -0.0587 0.5757 1.0000 4.000 0.7120 0.00771 0.00257 -0.0573 0.4660 1.0000 4.500 0.7588 0.00872 0.00311 -0.0560 0.3539 1.0000 5.000 0.8068 0.00969 0.00375 -0.0549 0.2693 1.0000 5.500 0.8550 0.01065 0.00443 -0.0539 0.1962 1.0000 6.000 0.9002 0.01198 0.00531 -0.0526 0.1059 1.0000 6.500 0.9453 0.01331 0.00643 -0.0512 0.0580 1.0000 7.000 0.9923 0.01439 0.00751 -0.0499 0.0410 1.0000 7.500 1.0369 0.01572 0.00885 -0.0483 0.0196 1.0000 8.000 1.0767 0.01764 0.01091 -0.0458 0.0128 1.0000 8.500 1.1120 0.01994 0.01344 -0.0427 0.0103 1.0000 9.000 1.1345 0.02437 0.01831 -0.0380 0.0092 1.0000 9.500 1.1696 0.02658 0.02082 -0.0353 0.0083 1.0000 10.000 1.1933 0.03053 0.02522 -0.0314 0.0078 1.0000 10.500 1.2019 0.03519 0.03039 -0.0259 0.0074 1.0000 11.000 1.1930 0.04041 0.03611 -0.0196 0.0071 1.0000 11.500 1.1722 0.04677 0.04295 -0.0151 0.0070 1.0000 12.000 1.1380 0.05565 0.05232 -0.0138 0.0071 1.0000 12.500 1.0982 0.06717 0.06426 -0.0174 0.0071 1.0000 13.000 1.0560 0.08245 0.07991 -0.0266 0.0072 1.0000 13.500 1.0101 0.10344 0.10119 -0.0410 0.0077 1.0000 14.000 0.9432 0.13524 0.13316 -0.0592 0.0086 1.0000