XFOIL Version 6.96 Calculated polar for: AG13 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.250 -0.5591 0.09934 0.09715 0.0180 1.0000 0.0125 -8.000 -0.5546 0.09560 0.09343 0.0160 1.0000 0.0128 -7.750 -0.5501 0.09180 0.08965 0.0137 1.0000 0.0131 -7.500 -0.5454 0.08793 0.08580 0.0111 1.0000 0.0135 -7.250 -0.5386 0.08374 0.08163 0.0073 1.0000 0.0139 -7.000 -0.5266 0.07886 0.07677 0.0015 1.0000 0.0145 -6.750 -0.5088 0.07337 0.07127 -0.0061 1.0000 0.0153 -6.500 -0.4842 0.06815 0.06603 -0.0139 1.0000 0.0160 -6.250 -0.4605 0.06211 0.05994 -0.0216 1.0000 0.0162 -6.000 -0.4356 0.05542 0.05314 -0.0288 1.0000 0.0162 -4.500 -0.2864 0.01856 0.01399 -0.0476 1.0000 0.0140 -4.250 -0.2589 0.01631 0.01137 -0.0475 1.0000 0.0144 -4.000 -0.2318 0.01548 0.01033 -0.0471 1.0000 0.0157 -3.750 -0.2044 0.01304 0.00756 -0.0468 1.0000 0.0151 -3.500 -0.1773 0.01150 0.00582 -0.0464 1.0000 0.0150 -3.250 -0.1503 0.01043 0.00461 -0.0459 1.0000 0.0155 -3.000 -0.1232 0.00955 0.00363 -0.0455 1.0000 0.0168 -2.750 -0.0958 0.00865 0.00265 -0.0453 1.0000 0.0223 -2.500 -0.0686 0.00818 0.00215 -0.0450 1.0000 0.0368 -2.250 -0.0414 0.00762 0.00180 -0.0449 1.0000 0.0888 -2.000 -0.0145 0.00720 0.00166 -0.0449 1.0000 0.1598 -1.750 0.0125 0.00681 0.00157 -0.0449 1.0000 0.2467 -1.500 0.0396 0.00647 0.00152 -0.0450 1.0000 0.3382 -1.250 0.0792 0.00607 0.00145 -0.0477 0.9935 0.4486 -1.000 0.1177 0.00567 0.00141 -0.0501 0.9832 0.5628 -0.750 0.1529 0.00511 0.00137 -0.0516 0.9687 0.7193 -0.500 0.1843 0.00444 0.00127 -0.0514 0.9463 1.0000 -0.250 0.2142 0.00446 0.00117 -0.0514 0.9145 1.0000 0.000 0.2398 0.00454 0.00110 -0.0504 0.8754 1.0000 0.250 0.2648 0.00469 0.00105 -0.0494 0.8337 1.0000 0.500 0.2904 0.00487 0.00102 -0.0485 0.7918 1.0000 0.750 0.3165 0.00506 0.00102 -0.0479 0.7500 1.0000 1.000 0.3429 0.00527 0.00102 -0.0474 0.7085 1.0000 1.250 0.3696 0.00549 0.00105 -0.0469 0.6677 1.0000 1.500 0.3965 0.00570 0.00109 -0.0466 0.6280 1.0000 1.750 0.4235 0.00592 0.00115 -0.0463 0.5896 1.0000 2.000 0.4506 0.00615 0.00122 -0.0460 0.5519 1.0000 2.250 0.4776 0.00639 0.00130 -0.0458 0.5135 1.0000 2.500 0.5047 0.00663 0.00140 -0.0456 0.4763 1.0000 2.750 0.5319 0.00688 0.00151 -0.0454 0.4398 1.0000 3.000 0.5589 0.00715 0.00166 -0.0453 0.4033 1.0000 3.250 0.5859 0.00744 0.00180 -0.0451 0.3665 1.0000 3.500 0.6130 0.00772 0.00197 -0.0450 0.3320 1.0000 3.750 0.6399 0.00804 0.00215 -0.0448 0.2971 1.0000 4.000 0.6668 0.00836 0.00238 -0.0447 0.2644 1.0000 4.250 0.6937 0.00868 0.00261 -0.0445 0.2342 1.0000 4.500 0.7203 0.00906 0.00286 -0.0444 0.2029 1.0000 4.750 0.7470 0.00942 0.00314 -0.0442 0.1756 1.0000 5.000 0.7735 0.00983 0.00344 -0.0440 0.1478 1.0000 5.250 0.7999 0.01024 0.00379 -0.0438 0.1236 1.0000 5.500 0.8263 0.01067 0.00416 -0.0436 0.1013 1.0000 5.750 0.8523 0.01116 0.00458 -0.0434 0.0820 1.0000 6.000 0.8785 0.01161 0.00501 -0.0432 0.0660 1.0000 6.250 0.9044 0.01212 0.00548 -0.0429 0.0513 1.0000 6.500 0.9301 0.01269 0.00601 -0.0426 0.0385 1.0000 6.750 0.9555 0.01334 0.00671 -0.0423 0.0283 1.0000 7.000 0.9804 0.01410 0.00751 -0.0418 0.0201 1.0000 7.250 1.0036 0.01538 0.00889 -0.0411 0.0144 1.0000 7.500 1.0283 0.01619 0.00982 -0.0406 0.0122 1.0000 7.750 1.0518 0.01725 0.01098 -0.0400 0.0104 1.0000 8.000 1.0692 0.01996 0.01398 -0.0386 0.0091 1.0000 8.250 1.0886 0.02214 0.01643 -0.0374 0.0086 1.0000 8.500 1.1118 0.02305 0.01749 -0.0368 0.0081 1.0000 8.750 1.1334 0.02432 0.01898 -0.0361 0.0074 1.0000 9.000 1.1519 0.02638 0.02130 -0.0351 0.0071 1.0000 9.250 1.1689 0.02863 0.02383 -0.0340 0.0068 1.0000 9.500 1.1840 0.03104 0.02653 -0.0330 0.0064 1.0000 9.750 1.1969 0.03355 0.02932 -0.0319 0.0061 1.0000 10.000 1.2066 0.03634 0.03240 -0.0308 0.0059 1.0000 10.250 1.2109 0.03967 0.03603 -0.0296 0.0057 1.0000 10.500 1.2080 0.04359 0.04027 -0.0284 0.0056 1.0000 10.750 1.1964 0.04770 0.04468 -0.0271 0.0056 1.0000 11.000 1.1787 0.05224 0.04946 -0.0274 0.0056 1.0000 11.250 1.1568 0.05948 0.05698 -0.0326 0.0058 1.0000