XFOIL Version 6.96 Calculated polar for: NASA SC(2)-0406 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.750 -0.6687 0.09117 0.08771 0.0044 1.0000 0.0430 -8.500 -0.6722 0.08464 0.08124 -0.0045 1.0000 0.0438 -8.250 -0.6721 0.07732 0.07374 -0.0162 1.0000 0.0445 -8.000 -0.6683 0.07254 0.06870 -0.0209 1.0000 0.0449 -7.750 -0.6640 0.06670 0.06263 -0.0241 1.0000 0.0452 -7.500 -0.6584 0.05942 0.05549 -0.0250 1.0000 0.0468 -7.250 -0.6449 0.05599 0.05204 -0.0255 1.0000 0.0484 -7.000 -0.6295 0.05257 0.04850 -0.0267 1.0000 0.0519 -6.750 -0.6116 0.04699 0.04220 -0.0304 1.0000 0.0590 -6.500 -0.5935 0.04358 0.03891 -0.0307 1.0000 0.0617 -6.250 -0.5715 0.04032 0.03542 -0.0316 1.0000 0.0657 -6.000 -0.5239 0.02179 0.01685 -0.0324 1.0000 0.0740 -5.750 -0.5024 0.01895 0.01383 -0.0327 1.0000 0.0772 -5.500 -0.4828 0.02282 0.01555 -0.0324 1.0000 0.0318 -5.250 -0.4548 0.02036 0.01270 -0.0323 1.0000 0.0321 -5.000 -0.4268 0.01881 0.01086 -0.0321 1.0000 0.0330 -4.750 -0.4006 0.01652 0.00853 -0.0322 1.0000 0.0382 -4.500 -0.3730 0.01529 0.00720 -0.0319 1.0000 0.0428 -4.250 -0.3454 0.01375 0.00568 -0.0319 1.0000 0.0527 -4.000 -0.3174 0.01295 0.00481 -0.0321 1.0000 0.0715 -3.750 -0.2894 0.01212 0.00410 -0.0325 1.0000 0.0951 -3.500 -0.2603 0.01115 0.00332 -0.0332 1.0000 0.1330 -3.250 -0.2306 0.00870 0.00282 -0.0353 1.0000 0.5682 -3.000 -0.2061 0.00845 0.00287 -0.0341 1.0000 0.6806 -2.750 -0.1816 0.00839 0.00288 -0.0329 1.0000 0.7275 -2.500 -0.1576 0.00837 0.00290 -0.0317 1.0000 0.7632 -2.250 -0.1343 0.00837 0.00292 -0.0302 1.0000 0.7934 -2.000 -0.1127 0.00836 0.00295 -0.0283 1.0000 0.8187 -1.750 -0.0932 0.00836 0.00298 -0.0259 1.0000 0.8442 -1.500 -0.0775 0.00835 0.00304 -0.0225 1.0000 0.8732 -1.250 -0.0641 0.00826 0.00301 -0.0187 1.0000 0.8994 -1.000 -0.0479 0.00811 0.00289 -0.0158 1.0000 0.9209 -0.750 -0.0301 0.00792 0.00272 -0.0135 1.0000 0.9417 -0.500 -0.0073 0.00772 0.00254 -0.0124 1.0000 0.9648 -0.250 0.0227 0.00758 0.00243 -0.0133 1.0000 0.9994 0.000 0.0545 0.00761 0.00246 -0.0150 1.0000 1.0000 0.250 0.0864 0.00767 0.00253 -0.0166 1.0000 1.0000 0.500 0.1180 0.00776 0.00265 -0.0181 1.0000 1.0000 0.750 0.1489 0.00787 0.00281 -0.0194 1.0000 1.0000 1.000 0.1793 0.00801 0.00300 -0.0205 1.0000 1.0000 1.250 0.2134 0.00813 0.00321 -0.0224 0.9973 1.0000 1.500 0.2797 0.00789 0.00317 -0.0300 0.9714 1.0000 1.750 0.3296 0.00766 0.00305 -0.0337 0.9262 1.0000 2.000 0.3623 0.00762 0.00298 -0.0334 0.8460 1.0000 2.250 0.3795 0.00847 0.00289 -0.0296 0.5911 1.0000 2.500 0.3963 0.01132 0.00354 -0.0287 0.1375 1.0000 2.750 0.4236 0.01237 0.00434 -0.0288 0.0950 1.0000 3.000 0.4517 0.01304 0.00494 -0.0291 0.0710 1.0000 3.250 0.4788 0.01409 0.00594 -0.0291 0.0515 1.0000 3.500 0.5052 0.01576 0.00759 -0.0287 0.0409 1.0000 3.750 0.5335 0.01702 0.00900 -0.0284 0.0375 1.0000 4.000 0.5617 0.01874 0.01089 -0.0280 0.0357 1.0000 4.250 0.5897 0.02028 0.01261 -0.0278 0.0333 1.0000 4.500 0.6158 0.02273 0.01522 -0.0278 0.0304 1.0000 4.750 0.6428 0.02536 0.01824 -0.0273 0.0306 1.0000 5.000 0.6699 0.03099 0.02491 -0.0254 0.0378 1.0000 5.250 0.6967 0.04002 0.03457 -0.0235 0.0683 1.0000 5.500 0.7176 0.04322 0.03856 -0.0230 0.0595 1.0000 5.750 0.7371 0.04647 0.04195 -0.0230 0.0569 1.0000 6.000 0.7546 0.05041 0.04587 -0.0231 0.0552 1.0000 6.250 0.7629 0.05835 0.05399 -0.0239 0.0537 1.0000 6.500 0.7815 0.06109 0.05751 -0.0250 0.0481 1.0000 6.750 0.7932 0.06545 0.06203 -0.0261 0.0456 1.0000 7.000 0.8045 0.06830 0.06488 -0.0258 0.0433 1.0000 7.250 0.8041 0.07651 0.07289 -0.0259 0.0413 1.0000 7.500 0.8075 0.08124 0.07784 -0.0274 0.0412 1.0000 7.750 0.8094 0.08610 0.08291 -0.0300 0.0410 1.0000 8.000 0.8077 0.09185 0.08882 -0.0356 0.0405 1.0000 8.250 0.8024 0.09805 0.09506 -0.0431 0.0402 1.0000 8.500 0.7984 0.10418 0.10116 -0.0506 0.0395 1.0000