XFOIL Version 6.94 Calculated polar for: GOE 795 smoothed 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 0.3805 0.00723 0.00276 -0.0851 0.8964 1.0000 0.500 0.4413 0.00716 0.00253 -0.0858 0.8698 1.0000 1.000 0.4919 0.00721 0.00246 -0.0843 0.8397 1.0000 1.500 0.5403 0.00732 0.00246 -0.0823 0.8092 1.0000 2.000 0.5875 0.00747 0.00253 -0.0800 0.7765 1.0000 2.500 0.6322 0.00766 0.00259 -0.0772 0.7346 1.0000 3.000 0.6763 0.00790 0.00271 -0.0742 0.6891 1.0000 3.500 0.7196 0.00821 0.00293 -0.0711 0.6385 1.0000 4.000 0.7612 0.00864 0.00319 -0.0678 0.5770 1.0000 4.500 0.7997 0.00928 0.00358 -0.0640 0.4967 1.0000 5.000 0.8347 0.01020 0.00411 -0.0597 0.3970 1.0000 5.500 0.8673 0.01143 0.00482 -0.0553 0.2793 1.0000 6.000 0.8986 0.01298 0.00580 -0.0510 0.1577 1.0000 6.500 0.9318 0.01453 0.00696 -0.0470 0.0806 1.0000 7.000 0.9608 0.01654 0.00876 -0.0420 0.0366 1.0000 8.000 1.0142 0.02124 0.01379 -0.0316 0.0227 1.0000 8.500 1.0464 0.02424 0.01708 -0.0274 0.0209 1.0000 9.000 1.0803 0.02648 0.01958 -0.0239 0.0180 1.0000 9.500 1.1104 0.02999 0.02344 -0.0200 0.0167 1.0000 10.000 1.1310 0.03439 0.02825 -0.0154 0.0156 1.0000 10.500 1.1366 0.03986 0.03433 -0.0090 0.0156 1.0000 11.000 1.1189 0.04599 0.04114 -0.0004 0.0160 1.0000 11.500 1.0609 0.05631 0.05227 0.0072 0.0177 1.0000 12.000 1.0114 0.06659 0.06301 0.0075 0.0183 1.0000 12.500 0.9633 0.07963 0.07639 0.0013 0.0188 1.0000 13.000 0.9153 0.09766 0.09468 -0.0113 0.0190 1.0000 13.500 0.8567 0.12539 0.12245 -0.0269 0.0207 1.0000