XFOIL Version 6.94 Calculated polar for: GOA 2 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1610 0.02738 0.01444 -0.0322 1.0000 0.1451 -2.750 -0.1342 0.02570 0.01273 -0.0308 1.0000 0.1628 -2.500 -0.1082 0.02372 0.01118 -0.0298 1.0000 0.2130 -2.250 -0.0859 0.01902 0.00944 -0.0262 1.0000 1.0000 -2.000 -0.0603 0.01903 0.00872 -0.0248 1.0000 1.0000 -1.750 -0.0355 0.01908 0.00835 -0.0237 1.0000 1.0000 -1.500 -0.0110 0.01913 0.00809 -0.0226 1.0000 1.0000 -1.250 0.0135 0.01920 0.00789 -0.0216 1.0000 1.0000 -1.000 0.0380 0.01927 0.00778 -0.0206 1.0000 1.0000 -0.750 0.0628 0.01936 0.00780 -0.0196 1.0000 1.0000 -0.500 0.0882 0.01945 0.00796 -0.0185 1.0000 1.0000 -0.250 0.1156 0.01950 0.00838 -0.0174 1.0000 1.0000 0.000 0.2035 0.02284 0.00890 -0.0249 0.3136 1.0000 0.250 0.2278 0.02409 0.00964 -0.0237 0.2924 1.0000 0.500 0.2543 0.02524 0.01047 -0.0229 0.2741 1.0000 0.750 0.2822 0.02641 0.01136 -0.0223 0.2562 1.0000 1.000 0.3118 0.02770 0.01244 -0.0220 0.2433 1.0000 1.250 0.3444 0.02888 0.01370 -0.0220 0.2398 1.0000 1.500 0.3777 0.03019 0.01518 -0.0222 0.2382 1.0000 1.750 0.4113 0.03163 0.01686 -0.0225 0.2383 1.0000 2.000 0.4436 0.03324 0.01873 -0.0227 0.2399 1.0000 2.250 0.4750 0.03511 0.02085 -0.0229 0.2424 1.0000 2.500 0.5073 0.03681 0.02313 -0.0234 0.2485 1.0000 2.750 0.5379 0.03907 0.02577 -0.0240 0.2556 1.0000 3.000 0.5695 0.04138 0.02864 -0.0250 0.2657 1.0000 3.250 0.5986 0.04439 0.03186 -0.0257 0.2757 1.0000 3.500 0.6324 0.04713 0.03548 -0.0287 0.2981 1.0000 3.750 0.6640 0.05076 0.03962 -0.0318 0.3239 1.0000 4.000 0.6988 0.05483 0.04447 -0.0391 0.3669 1.0000 4.250 0.7303 0.06007 0.05031 -0.0498 0.4235 1.0000 4.750 0.6736 0.07464 0.06506 -0.0871 0.6666 1.0000 5.000 0.6100 0.07783 0.06799 -0.0883 0.7683 1.0000