XFOIL Version 6.94 Calculated polar for: HQ 3.0/9 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 0.4126 0.00662 0.00214 -0.0908 0.8110 0.8038 0.500 0.4663 0.00655 0.00208 -0.0901 0.7925 0.8293 1.000 0.5185 0.00645 0.00201 -0.0890 0.7716 0.8618 1.500 0.5719 0.00624 0.00193 -0.0879 0.7482 0.9376 2.000 0.6354 0.00634 0.00194 -0.0897 0.7229 1.0000 2.500 0.6901 0.00653 0.00204 -0.0895 0.6936 1.0000 3.000 0.7436 0.00678 0.00218 -0.0890 0.6574 1.0000 3.500 0.7958 0.00712 0.00238 -0.0882 0.6111 1.0000 4.000 0.8457 0.00762 0.00270 -0.0871 0.5456 1.0000 4.500 0.8927 0.00841 0.00314 -0.0856 0.4594 1.0000 5.000 0.9331 0.00985 0.00383 -0.0833 0.3165 1.0000 5.500 0.9692 0.01191 0.00479 -0.0808 0.1443 1.0000 6.000 1.0052 0.01411 0.00620 -0.0782 0.0188 1.0000 6.500 1.0510 0.01523 0.00727 -0.0766 0.0028 1.0000 7.000 1.0977 0.01620 0.00845 -0.0751 0.0023 1.0000 7.500 1.1418 0.01743 0.00995 -0.0731 0.0023 1.0000 8.000 1.1815 0.01908 0.01189 -0.0705 0.0023 1.0000 8.500 1.2157 0.02111 0.01426 -0.0672 0.0024 1.0000 9.000 1.2426 0.02363 0.01707 -0.0629 0.0026 1.0000 9.500 1.2582 0.02678 0.02059 -0.0569 0.0029 1.0000 10.000 1.2670 0.03142 0.02570 -0.0506 0.0033 1.0000 10.500 1.2686 0.03764 0.03254 -0.0445 0.0039 1.0000 11.000 1.2544 0.04480 0.04029 -0.0388 0.0043 1.0000 11.500 1.2236 0.05356 0.04961 -0.0350 0.0047 1.0000 12.000 1.1864 0.06391 0.06042 -0.0348 0.0048 1.0000 12.500 1.1493 0.07596 0.07286 -0.0387 0.0049 1.0000 13.000 1.1194 0.08915 0.08638 -0.0461 0.0049 1.0000 13.500 1.0868 0.10608 0.10361 -0.0571 0.0048 1.0000 14.000 1.0628 0.12378 0.12154 -0.0687 0.0047 1.0000 14.500 1.0402 0.14299 0.14092 -0.0803 0.0045 1.0000 15.000 1.0145 0.16495 0.16292 -0.0919 0.0044 1.0000 15.500 1.0015 0.18306 0.18098 -0.1009 0.0047 1.0000