XFOIL Version 6.96 Calculated polar for: FX 60-177 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 ------ -------- --------- --------- -------- -------- -------- -13.500 -0.4669 0.10926 0.10606 -0.0512 1.0001 0.0249 -13.250 -0.4590 0.10983 0.10670 -0.0491 1.0001 0.0253 -13.000 -0.6353 0.08724 0.08332 -0.0677 1.0001 0.0238 -12.750 -0.6566 0.08273 0.07876 -0.0679 1.0001 0.0236 -12.500 -0.6803 0.07848 0.07443 -0.0677 1.0001 0.0234 -12.250 -0.7046 0.07460 0.07045 -0.0668 1.0001 0.0232 -12.000 -0.7291 0.07105 0.06681 -0.0653 1.0001 0.0229 -11.750 -0.7539 0.06779 0.06343 -0.0633 1.0001 0.0227 -11.500 -0.7791 0.06475 0.06026 -0.0606 1.0001 0.0224 -11.250 -0.8051 0.06188 0.05724 -0.0573 1.0001 0.0221 -11.000 -0.8789 0.05496 0.04941 -0.0517 0.9966 0.0199 -10.750 -0.8812 0.05307 0.04720 -0.0511 0.9918 0.0196 -10.500 -0.8735 0.04969 0.04361 -0.0509 0.9882 0.0195 -10.250 -0.8622 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-0.0710 0.6690 0.8180 4.500 0.7964 0.01761 0.01166 -0.0694 0.6476 0.8208 4.750 0.8128 0.01759 0.01159 -0.0670 0.6248 0.8235 5.000 0.8335 0.01765 0.01154 -0.0657 0.6011 0.8266 5.250 0.8548 0.01779 0.01160 -0.0649 0.5759 0.8297 5.500 0.8804 0.01803 0.01173 -0.0653 0.5501 0.8329 5.750 0.9013 0.01830 0.01188 -0.0645 0.5254 0.8358 6.000 0.9142 0.01858 0.01209 -0.0618 0.5016 0.8382 6.250 0.9301 0.01894 0.01235 -0.0601 0.4779 0.8407 6.500 0.9477 0.01935 0.01269 -0.0589 0.4540 0.8433 6.750 0.9670 0.01982 0.01307 -0.0582 0.4307 0.8461 7.000 0.9901 0.02039 0.01353 -0.0584 0.4093 0.8491 7.250 1.0047 0.02082 0.01392 -0.0566 0.3899 0.8516 7.500 1.0174 0.02126 0.01433 -0.0544 0.3713 0.8540 7.750 1.0327 0.02180 0.01480 -0.0529 0.3534 0.8565 8.000 1.0503 0.02238 0.01534 -0.0520 0.3366 0.8590 8.250 1.0700 0.02299 0.01593 -0.0516 0.3205 0.8615 8.500 1.0920 0.02368 0.01660 -0.0518 0.3054 0.8643 8.750 1.1067 0.02423 0.01713 -0.0502 0.2928 0.8667 9.000 1.1196 0.02484 0.01769 -0.0484 0.2809 0.8691 9.250 1.1344 0.02549 0.01835 -0.0470 0.2682 0.8716 9.500 1.1508 0.02616 0.01907 -0.0461 0.2557 0.8744 9.750 1.1685 0.02694 0.01986 -0.0456 0.2439 0.8773 10.000 1.1877 0.02785 0.02072 -0.0456 0.2333 0.8796 10.250 1.2013 0.02853 0.02147 -0.0442 0.2237 0.8817 10.500 1.2148 0.02928 0.02227 -0.0428 0.2148 0.8840 10.750 1.2269 0.03013 0.02312 -0.0414 0.2058 0.8867 11.000 1.2408 0.03099 0.02408 -0.0404 0.1959 0.8895 11.250 1.2542 0.03206 0.02515 -0.0397 0.1867 0.8921 11.500 1.2687 0.03321 0.02633 -0.0394 0.1773 0.8942 11.750 1.2795 0.03422 0.02742 -0.0381 0.1688 0.8963 12.000 1.2875 0.03538 0.02860 -0.0366 0.1608 0.8986 12.250 1.2985 0.03648 0.02983 -0.0356 0.1523 0.9011 12.500 1.3061 0.03794 0.03127 -0.0345 0.1446 0.9035 12.750 1.3179 0.03925 0.03271 -0.0340 0.1361 0.9062 13.000 1.3263 0.04099 0.03444 -0.0335 0.1294 0.9087 13.250 1.3343 0.04232 0.03591 -0.0323 0.1228 0.9114 13.500 1.3373 0.04413 0.03772 -0.0310 0.1172 0.9140 13.750 1.3464 0.04569 0.03946 -0.0303 0.1111 0.9166 14.000 1.3523 0.04766 0.04148 -0.0298 0.1055 0.9191 14.250 1.3584 0.04977 0.04366 -0.0296 0.1000 0.9216 14.500 1.3643 0.05193 0.04592 -0.0295 0.0942 0.9239 14.750 1.3628 0.05444 0.04842 -0.0286 0.0896 0.9264 15.000 1.3686 0.05647 0.05063 -0.0283 0.0846 0.9293 15.250 1.3697 0.05916 0.05338 -0.0283 0.0801 0.9324 15.500 1.3701 0.06209 0.05637 -0.0284 0.0764 0.9353 15.750 1.3741 0.06456 0.05901 -0.0285 0.0725 0.9383 16.000 1.3731 0.06744 0.06201 -0.0285 0.0693 0.9416 16.250 1.3692 0.07093 0.06554 -0.0289 0.0666 0.9446 16.500 1.3697 0.07409 0.06884 -0.0295 0.0641 0.9477 16.750 1.3697 0.07734 0.07225 -0.0303 0.0610 0.9509 17.000 1.3644 0.08097 0.07597 -0.0308 0.0585 0.9549 17.250 1.3574 0.08512 0.08020 -0.0318 0.0565 0.9589 17.500 1.3512 0.08923 0.08439 -0.0329 0.0547 0.9630 17.750 1.3480 0.09268 0.08803 -0.0337 0.0530 0.9685 18.000 1.3436 0.09661 0.09214 -0.0350 0.0509 0.9744 18.250 1.3359 0.10106 0.09672 -0.0368 0.0486 0.9831 18.500 1.3254 0.10631 0.10207 -0.0394 0.0464 0.9999