XFOIL Version 6.96 Calculated polar for: WORTMANN FX 08-S-176 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -11.000 -0.0150 0.09123 0.08724 -0.0847 0.5975 0.0199 -10.750 -0.0091 0.08840 0.08442 -0.0855 0.5968 0.0207 -10.500 -0.0066 0.08491 0.08095 -0.0868 0.5962 0.0211 -10.250 -0.0057 0.08116 0.07722 -0.0883 0.5955 0.0224 -10.000 -0.0060 0.07717 0.07326 -0.0899 0.5949 0.0228 -9.750 -0.0118 0.07219 0.06830 -0.0925 0.5944 0.0238 -9.500 -0.0204 0.06648 0.06263 -0.0964 0.5939 0.0244 -9.250 -0.0271 0.06113 0.05730 -0.0997 0.5933 0.0245 -9.000 -0.0425 0.05441 0.05059 -0.1052 0.5929 0.0245 -8.750 -0.0593 0.04972 0.04585 -0.1074 0.5923 0.0245 -8.500 -0.0774 0.04588 0.04196 -0.1074 0.5916 0.0245 -8.250 -0.0943 0.04296 0.03897 -0.1055 0.5909 0.0245 -8.000 -0.1113 0.03980 0.03569 -0.1027 0.5901 0.0245 -7.500 -0.1568 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-0.0697 0.5465 0.9181 2.250 0.6951 0.01387 0.00846 -0.0686 0.5452 0.9220 2.500 0.7216 0.01382 0.00838 -0.0686 0.5439 0.9241 2.750 0.7495 0.01379 0.00831 -0.0689 0.5428 0.9262 3.000 0.7783 0.01379 0.00826 -0.0694 0.5415 0.9281 3.250 0.8073 0.01394 0.00836 -0.0702 0.5402 0.9300 3.500 0.8332 0.01399 0.00847 -0.0703 0.5386 0.9326 3.750 0.8540 0.01396 0.00850 -0.0692 0.5368 0.9353 4.000 0.8775 0.01396 0.00854 -0.0687 0.5349 0.9377 4.250 0.9032 0.01396 0.00855 -0.0686 0.5332 0.9398 4.500 0.9298 0.01393 0.00855 -0.0688 0.5315 0.9419 4.750 0.9578 0.01390 0.00851 -0.0692 0.5301 0.9438 5.000 0.9872 0.01386 0.00846 -0.0700 0.5287 0.9454 5.250 1.0175 0.01385 0.00843 -0.0710 0.5274 0.9466 5.500 1.0440 0.01386 0.00843 -0.0710 0.5258 0.9484 5.750 1.0625 0.01385 0.00854 -0.0696 0.5234 0.9514 6.000 1.0855 0.01385 0.00861 -0.0692 0.5209 0.9536 6.250 1.1112 0.01381 0.00863 -0.0692 0.5186 0.9554 6.500 1.1390 0.01375 0.00859 -0.0697 0.5166 0.9567 6.750 1.1684 0.01369 0.00853 -0.0705 0.5148 0.9578 7.000 1.1997 0.01350 0.00831 -0.0715 0.5123 0.9588 7.250 1.2192 0.01338 0.00833 -0.0705 0.5076 0.9608 7.500 1.2440 0.01295 0.00791 -0.0702 0.5018 0.9623 7.750 1.2703 0.01271 0.00766 -0.0702 0.4973 0.9637 8.000 1.2902 0.01265 0.00773 -0.0693 0.4912 0.9657 8.250 1.3135 0.01242 0.00748 -0.0688 0.4837 0.9676 8.500 1.3311 0.01235 0.00747 -0.0674 0.4722 0.9706 8.750 1.3465 0.01238 0.00750 -0.0657 0.4572 0.9733 9.000 1.3568 0.01262 0.00772 -0.0632 0.4363 0.9773 9.250 1.3605 0.01330 0.00830 -0.0601 0.4096 0.9831 9.500 1.3575 0.01483 0.00968 -0.0575 0.3759 1.0000 9.750 1.3419 0.01721 0.01196 -0.0542 0.3506 1.0000 10.000 1.3323 0.02007 0.01481 -0.0529 0.3337 1.0000 10.250 1.3175 0.02371 0.01841 -0.0518 0.3160 1.0000 10.500 1.3028 0.02728 0.02195 -0.0505 0.2981 1.0000 10.750 1.2800 0.03136 0.02594 -0.0486 0.2724 1.0000 11.000 1.2647 0.03479 0.02930 -0.0470 0.2498 1.0000 11.250 1.2269 0.04009 0.03437 -0.0444 0.2109 1.0000 11.500 1.1980 0.04504 0.03910 -0.0426 0.1726 1.0000 11.750 1.1716 0.05000 0.04381 -0.0411 0.1275 1.0000 12.000 1.1527 0.05449 0.04808 -0.0402 0.0871 1.0000 12.250 1.1292 0.05967 0.05301 -0.0393 0.0415 1.0000 12.500 1.1233 0.06325 0.05650 -0.0392 0.0183 1.0000 12.750 1.1247 0.06612 0.05936 -0.0392 0.0130 1.0000 13.000 1.1296 0.06866 0.06197 -0.0394 0.0119 1.0000 13.250 1.1346 0.07122 0.06461 -0.0396 0.0109 1.0000 13.500 1.1410 0.07367 0.06713 -0.0399 0.0105 1.0000 13.750 1.1451 0.07643 0.06998 -0.0402 0.0100 1.0000 14.000 1.1504 0.07904 0.07266 -0.0406 0.0097 1.0000 14.250 1.1533 0.08201 0.07571 -0.0410 0.0094 1.0000 14.500 1.1559 0.08505 0.07883 -0.0414 0.0092 1.0000 14.750 1.1581 0.08820 0.08206 -0.0420 0.0090 1.0000 15.000 1.1580 0.09170 0.08563 -0.0426 0.0088 1.0000 15.250 1.1572 0.09532 0.08934 -0.0433 0.0086 1.0000 15.500 1.1559 0.09905 0.09316 -0.0440 0.0084 1.0000 15.750 1.1568 0.10254 0.09672 -0.0448 0.0084 1.0000