XFOIL Version 6.96 Calculated polar for: WORTMANN FX 05-191 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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -15.500 -0.4523 0.09241 0.08819 -0.1078 0.9495 0.0234 -15.250 -0.5065 0.07675 0.07216 -0.1194 0.9461 0.0230 -15.000 -0.5334 0.06787 0.06295 -0.1263 0.9417 0.0230 -14.750 -0.5490 0.06138 0.05614 -0.1314 0.9373 0.0231 -14.500 -0.5582 0.05613 0.05058 -0.1355 0.9332 0.0233 -14.250 -0.5644 0.05203 0.04619 -0.1380 0.9284 0.0235 -14.000 -0.5688 0.04847 0.04233 -0.1397 0.9234 0.0239 -13.750 -0.5698 0.04526 0.03877 -0.1411 0.9192 0.0242 -13.500 -0.5684 0.04275 0.03597 -0.1416 0.9148 0.0246 -13.250 -0.5636 0.04069 0.03367 -0.1416 0.9101 0.0249 -13.000 -0.5493 0.03904 0.03194 -0.1420 0.9066 0.0252 -12.750 -0.5343 0.03759 0.03038 -0.1424 0.9038 0.0256 -12.500 -0.5207 0.03629 0.02897 -0.1423 0.9009 0.0260 -12.250 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0.01071 -0.0754 0.5582 0.8109 7.000 1.0137 0.01756 0.01101 -0.0740 0.5443 0.8148 7.250 1.0286 0.01787 0.01133 -0.0721 0.5313 0.8177 7.500 1.0411 0.01821 0.01168 -0.0698 0.5184 0.8202 7.750 1.0520 0.01866 0.01217 -0.0674 0.5044 0.8230 8.000 1.0647 0.01914 0.01269 -0.0655 0.4899 0.8262 8.250 1.0788 0.01967 0.01324 -0.0640 0.4741 0.8294 8.500 1.0916 0.02028 0.01384 -0.0624 0.4559 0.8322 8.750 1.0981 0.02098 0.01453 -0.0596 0.4342 0.8345 9.000 1.1010 0.02192 0.01539 -0.0566 0.4053 0.8371 9.250 1.0996 0.02320 0.01651 -0.0532 0.3664 0.8398 9.500 1.0922 0.02496 0.01802 -0.0495 0.3216 0.8427 9.750 1.0838 0.02702 0.01983 -0.0461 0.2757 0.8456 10.000 1.0746 0.02923 0.02179 -0.0428 0.2317 0.8478 10.250 1.0673 0.03136 0.02374 -0.0398 0.1942 0.8497 10.500 1.0644 0.03339 0.02563 -0.0374 0.1637 0.8518 10.750 1.0639 0.03539 0.02752 -0.0355 0.1377 0.8540 11.000 1.0671 0.03728 0.02933 -0.0341 0.1171 0.8560 11.250 1.0718 0.03915 0.03116 -0.0331 0.1032 0.8578 11.500 1.0786 0.04097 0.03296 -0.0323 0.0939 0.8598 11.750 1.0859 0.04283 0.03483 -0.0317 0.0870 0.8617 12.000 1.0933 0.04452 0.03657 -0.0309 0.0817 0.8634 12.250 1.1001 0.04629 0.03840 -0.0301 0.0769 0.8649 12.500 1.1053 0.04832 0.04046 -0.0293 0.0729 0.8664 12.750 1.1153 0.04997 0.04222 -0.0289 0.0690 0.8680 13.000 1.1230 0.05187 0.04419 -0.0285 0.0650 0.8697 13.250 1.1279 0.05412 0.04645 -0.0281 0.0617 0.8716 13.500 1.1386 0.05586 0.04832 -0.0280 0.0583 0.8735 13.750 1.1473 0.05788 0.05041 -0.0280 0.0550 0.8752 14.000 1.1531 0.06027 0.05282 -0.0280 0.0523 0.8766 14.250 1.1626 0.06228 0.05494 -0.0281 0.0496 0.8780 14.500 1.1704 0.06432 0.05708 -0.0280 0.0470 0.8793 14.750 1.1764 0.06662 0.05942 -0.0280 0.0447 0.8807 15.000 1.1819 0.06902 0.06187 -0.0280 0.0428 0.8822 15.250 1.1895 0.07121 0.06418 -0.0281 0.0408 0.8840 15.500 1.1965 0.07354 0.06660 -0.0283 0.0389 0.8859 15.750 1.2018 0.07613 0.06924 -0.0287 0.0374 0.8877 16.000 1.2070 0.07878 0.07195 -0.0291 0.0359 0.8893 16.250 1.2145 0.08122 0.07451 -0.0296 0.0344 0.8909 16.500 1.2209 0.08383 0.07721 -0.0302 0.0330 0.8924 17.000 1.2288 0.08958 0.08309 -0.0315 0.0308 0.8953 17.250 1.2336 0.09230 0.08594 -0.0320 0.0297 0.8969 17.500 1.2386 0.09504 0.08880 -0.0326 0.0287 0.8986 17.750 1.2428 0.09794 0.09182 -0.0334 0.0277 0.9004 18.000 1.2463 0.10102 0.09498 -0.0344 0.0268 0.9023 18.250 1.2487 0.10428 0.09832 -0.0355 0.0260 0.9043 18.500 1.2506 0.10763 0.10172 -0.0367 0.0253 0.9064 18.750 1.2545 0.11078 0.10504 -0.0379 0.0245 0.9086 19.000 1.2562 0.11399 0.10840 -0.0390 0.0237 0.9110 19.250 1.2576 0.11731 0.11184 -0.0402 0.0231 0.9136