XFOIL Version 6.94 Calculated polar for: FX 84-W-097 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.2896 0.00785 0.00282 -0.0546 0.7190 1.0000 0.500 0.3406 0.00788 0.00277 -0.0540 0.7035 1.0000 1.000 0.3917 0.00805 0.00275 -0.0531 0.6885 1.0000 1.500 0.4439 0.00813 0.00276 -0.0525 0.6720 1.0000 2.000 0.4969 0.00833 0.00285 -0.0519 0.6558 1.0000 2.500 0.5509 0.00847 0.00294 -0.0515 0.6388 1.0000 3.000 0.6051 0.00869 0.00316 -0.0512 0.6211 1.0000 3.500 0.6593 0.00889 0.00328 -0.0508 0.6015 1.0000 4.000 0.7139 0.00909 0.00355 -0.0506 0.5822 1.0000 4.500 0.7680 0.00937 0.00380 -0.0502 0.5604 1.0000 5.000 0.8210 0.00954 0.00403 -0.0496 0.5248 1.0000 5.500 0.8736 0.00984 0.00431 -0.0490 0.4849 1.0000 6.000 0.9243 0.01037 0.00471 -0.0482 0.4311 1.0000 6.500 0.9706 0.01138 0.00537 -0.0471 0.3364 1.0000 7.000 1.0049 0.01393 0.00700 -0.0451 0.1545 1.0000 7.500 1.0345 0.01711 0.00950 -0.0425 0.0473 1.0000 9.000 1.1279 0.02430 0.01707 -0.0341 0.0225 1.0000 9.500 1.1512 0.02718 0.02012 -0.0305 0.0200 1.0000 10.000 1.1774 0.03057 0.02376 -0.0275 0.0188 1.0000 10.500 1.2060 0.03786 0.03152 -0.0250 0.0169 1.0000 11.000 1.2197 0.04032 0.03434 -0.0216 0.0157 1.0000 11.500 1.2259 0.04536 0.03982 -0.0187 0.0152 1.0000 12.000 1.2139 0.05277 0.04776 -0.0162 0.0157 1.0000 12.500 1.1911 0.06099 0.05646 -0.0155 0.0158 1.0000 13.000 1.1686 0.06914 0.06500 -0.0170 0.0152 1.0000 13.500 1.1392 0.07951 0.07575 -0.0208 0.0151 1.0000 14.000 1.0943 0.09404 0.09066 -0.0282 0.0156 1.0000 14.500 1.0427 0.11307 0.11003 -0.0401 0.0159 1.0000