XFOIL Version 6.96 Calculated polar for: WORTMANN FX 66-17AII-182 AIRFOIL (AS TESTED AT N 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 ------ -------- --------- --------- -------- -------- -------- -5.000 -0.1403 0.01352 0.00557 -0.0810 0.5609 0.0417 -4.500 -0.0911 0.01190 0.00453 -0.0813 0.5568 0.1767 -4.000 -0.0427 0.01033 0.00390 -0.0816 0.5528 0.4320 -3.500 0.0137 0.01025 0.00391 -0.0821 0.5486 0.5024 -3.000 0.0713 0.01030 0.00400 -0.0825 0.5448 0.5413 -2.500 0.1294 0.01051 0.00414 -0.0831 0.5409 0.5679 -2.000 0.1875 0.01076 0.00430 -0.0836 0.5374 0.5879 -1.500 0.2458 0.01093 0.00439 -0.0843 0.5338 0.5980 -1.000 0.3040 0.01113 0.00452 -0.0850 0.5303 0.6102 -0.500 0.3620 0.01118 0.00459 -0.0857 0.5269 0.6188 0.000 0.4205 0.01131 0.00464 -0.0865 0.5235 0.6273 0.500 0.4786 0.01136 0.00469 -0.0873 0.5203 0.6346 1.000 0.5369 0.01154 0.00481 -0.0882 0.5172 0.6426 1.500 0.5946 0.01178 0.00505 -0.0890 0.5137 0.6500 2.000 0.6521 0.01189 0.00522 -0.0897 0.5108 0.6573 2.500 0.7098 0.01205 0.00540 -0.0906 0.5076 0.6646 3.000 0.7671 0.01217 0.00559 -0.0913 0.5045 0.6719 3.500 0.8249 0.01234 0.00575 -0.0922 0.5013 0.6793 4.000 0.8824 0.01268 0.00611 -0.0930 0.4980 0.6863 4.500 0.9385 0.01294 0.00650 -0.0937 0.4953 0.6940 5.000 0.9946 0.01313 0.00681 -0.0943 0.4919 0.7013 5.500 1.0506 0.01332 0.00713 -0.0949 0.4886 0.7092 6.000 1.1073 0.01356 0.00744 -0.0957 0.4854 0.7176 6.500 1.1638 0.01381 0.00778 -0.0965 0.4820 0.7262 7.500 1.2714 0.01377 0.00805 -0.0968 0.4688 0.7464 8.000 1.3240 0.01366 0.00806 -0.0967 0.4595 0.7580 9.000 1.4253 0.01358 0.00837 -0.0960 0.4391 0.7882 9.500 1.4717 0.01356 0.00856 -0.0949 0.4203 0.8106 10.000 1.5064 0.01389 0.00906 -0.0919 0.3818 0.8684