XFOIL Version 6.94 Calculated polar for: SA7035 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3091 0.00574 0.00175 -0.0620 0.7765 0.9783 0.500 0.3856 0.00581 0.00160 -0.0663 0.7397 1.0000 1.000 0.4336 0.00602 0.00157 -0.0646 0.7018 1.0000 1.500 0.4835 0.00627 0.00161 -0.0633 0.6627 1.0000 2.000 0.5341 0.00657 0.00170 -0.0622 0.6203 1.0000 2.500 0.5855 0.00690 0.00183 -0.0612 0.5776 1.0000 3.000 0.6373 0.00725 0.00205 -0.0604 0.5364 1.0000 3.500 0.6890 0.00766 0.00229 -0.0596 0.4935 1.0000 4.000 0.7403 0.00812 0.00261 -0.0588 0.4486 1.0000 4.500 0.7895 0.00878 0.00295 -0.0578 0.3766 1.0000 5.000 0.8379 0.00957 0.00338 -0.0567 0.3010 1.0000 5.500 0.8862 0.01043 0.00392 -0.0557 0.2341 1.0000 6.000 0.9305 0.01174 0.00473 -0.0543 0.1382 1.0000 6.500 0.9769 0.01282 0.00558 -0.0531 0.0862 1.0000 7.000 1.0184 0.01443 0.00685 -0.0512 0.0251 1.0000 7.500 1.0613 0.01583 0.00821 -0.0492 0.0057 1.0000 8.000 1.1059 0.01699 0.00960 -0.0474 0.0050 1.0000 8.500 1.1460 0.01852 0.01147 -0.0450 0.0048 1.0000 9.000 1.1793 0.02050 0.01374 -0.0417 0.0048 1.0000 9.500 1.2046 0.02287 0.01639 -0.0374 0.0048 1.0000 10.000 1.2193 0.02540 0.01920 -0.0317 0.0049 1.0000 10.500 1.2288 0.02831 0.02240 -0.0260 0.0051 1.0000 11.000 1.2340 0.03211 0.02655 -0.0209 0.0053 1.0000 11.500 1.2329 0.03725 0.03213 -0.0165 0.0056 1.0000 12.000 1.2212 0.04423 0.03964 -0.0133 0.0061 1.0000 12.500 1.1968 0.05313 0.04907 -0.0124 0.0065 1.0000 13.000 1.1675 0.06351 0.05989 -0.0148 0.0068 1.0000 13.500 1.1337 0.07614 0.07291 -0.0207 0.0070 1.0000 14.000 1.0986 0.09101 0.08812 -0.0294 0.0071 1.0000 14.500 1.0666 0.10741 0.10481 -0.0398 0.0071 1.0000 15.000 1.0314 0.12684 0.12448 -0.0521 0.0071 1.0000 15.500 0.9922 0.14993 0.14776 -0.0657 0.0071 1.0000 16.000 0.9511 0.17732 0.17518 -0.0793 0.0076 1.0000