XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY RC08-64C 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 ------ -------- --------- --------- -------- -------- -------- -8.250 -0.5632 0.08745 0.08382 -0.0172 1.0000 0.0374 -8.000 -0.5717 0.08191 0.07829 -0.0235 1.0000 0.0378 -7.750 -0.5818 0.07572 0.07195 -0.0297 1.0000 0.0383 -7.250 -0.5719 0.06920 0.06548 -0.0282 1.0000 0.0400 -7.000 -0.5669 0.06590 0.06212 -0.0284 1.0000 0.0416 -6.750 -0.5756 0.06137 0.05690 -0.0294 1.0000 0.0447 -6.500 -0.5621 0.05728 0.05315 -0.0287 1.0000 0.0458 -6.250 -0.5529 0.05499 0.05088 -0.0273 1.0000 0.0477 -6.000 -0.5516 0.05127 0.04661 -0.0255 1.0000 0.0526 -5.750 -0.5087 0.03469 0.03076 -0.0240 1.0000 0.0555 -5.500 -0.5122 0.03181 0.02732 -0.0206 1.0000 0.0613 -5.250 -0.5216 0.04317 0.03835 -0.0204 1.0000 0.0630 -5.000 -0.5087 0.04139 0.03647 -0.0186 1.0000 0.0681 -4.750 -0.4989 0.03870 0.03360 -0.0165 1.0000 0.0743 -4.500 -0.4873 0.03684 0.03147 -0.0143 1.0000 0.0847 -4.250 -0.4738 0.03592 0.03025 -0.0120 1.0000 0.0967 -4.000 -0.4585 0.03288 0.02741 -0.0110 1.0000 0.1007 -3.750 -0.4434 0.03113 0.02544 -0.0091 1.0000 0.1131 -3.500 -0.4270 0.02965 0.02377 -0.0075 1.0000 0.1273 -3.250 -0.4083 0.02792 0.02204 -0.0062 1.0000 0.1357 -3.000 -0.3899 0.02624 0.02027 -0.0049 1.0000 0.1509 -2.750 -0.3353 0.01845 0.01040 -0.0010 1.0000 0.0353 -2.500 -0.3096 0.01732 0.00905 0.0001 1.0000 0.0321 -2.250 -0.2841 0.01622 0.00791 0.0008 1.0000 0.0305 -2.000 -0.2595 0.01544 0.00707 0.0016 1.0000 0.0293 -1.750 -0.2359 0.01485 0.00646 0.0024 1.0000 0.0288 -1.500 -0.2125 0.01430 0.00594 0.0030 1.0000 0.0289 -1.250 -0.1790 0.01378 0.00544 0.0016 0.9975 0.0297 -1.000 -0.1383 0.01342 0.00505 -0.0012 0.9932 0.0336 -0.750 -0.0976 0.01308 0.00476 -0.0040 0.9883 0.0431 -0.500 -0.0752 0.01001 0.00480 -0.0034 0.9872 0.7909 -0.250 -0.0024 0.01026 0.00530 -0.0107 0.9940 1.0000 0.000 0.0429 0.01034 0.00526 -0.0146 0.9882 1.0000 0.250 0.0872 0.01040 0.00524 -0.0184 0.9818 1.0000 0.500 0.1343 0.01043 0.00524 -0.0227 0.9765 1.0000 0.750 0.1779 0.01042 0.00523 -0.0261 0.9686 1.0000 1.000 0.2349 0.01029 0.00514 -0.0322 0.9635 1.0000 1.250 0.3048 0.00949 0.00440 -0.0391 0.9362 1.0000 1.500 0.3367 0.00907 0.00399 -0.0386 0.9027 1.0000 1.750 0.3574 0.00878 0.00360 -0.0356 0.8480 1.0000 2.000 0.3782 0.00873 0.00334 -0.0330 0.7793 1.0000 2.250 0.3974 0.00903 0.00314 -0.0302 0.6619 1.0000 2.500 0.4085 0.01060 0.00320 -0.0269 0.3506 1.0000 2.750 0.4214 0.01277 0.00385 -0.0251 0.0438 1.0000 3.000 0.4434 0.01327 0.00431 -0.0239 0.0346 1.0000 3.500 0.4870 0.01421 0.00533 -0.0214 0.0286 1.0000 3.750 0.5079 0.01480 0.00596 -0.0200 0.0277 1.0000 4.000 0.5280 0.01550 0.00668 -0.0184 0.0272 1.0000 4.250 0.5478 0.01633 0.00749 -0.0168 0.0271 1.0000 4.500 0.5682 0.01727 0.00846 -0.0152 0.0273 1.0000 4.750 0.5897 0.01837 0.00959 -0.0137 0.0279 1.0000 5.000 0.6126 0.01969 0.01094 -0.0125 0.0288 1.0000 5.250 0.6369 0.02143 0.01264 -0.0116 0.0297 1.0000 5.500 0.6619 0.02292 0.01457 -0.0099 0.0329 1.0000 7.250 0.7380 0.03263 0.02794 0.0102 0.0793 1.0000 7.500 0.7934 0.05032 0.04546 0.0082 0.0703 1.0000 7.750 0.8076 0.05297 0.04809 0.0091 0.0674 1.0000 8.000 0.8220 0.05994 0.05477 0.0085 0.0653 1.0000 8.250 0.8084 0.06013 0.05582 0.0122 0.0594 1.0000 8.500 0.8209 0.06285 0.05849 0.0130 0.0574 1.0000 8.750 0.8329 0.06914 0.06458 0.0126 0.0559 1.0000 9.000 0.7944 0.07220 0.06833 0.0143 0.0541 1.0000 9.250 0.7747 0.07670 0.07296 0.0141 0.0532 1.0000 9.500 0.7551 0.08223 0.07857 0.0112 0.0528 1.0000