XFOIL Version 6.96 Calculated polar for: FX 63-147 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 ------ -------- --------- --------- -------- -------- -------- -9.250 -0.3122 0.08927 0.08648 -0.0395 1.0001 0.0499 -9.000 -0.4035 0.08515 0.08209 -0.0502 1.0001 0.0449 -8.750 -0.4315 0.07986 0.07684 -0.0525 1.0001 0.0447 -8.000 -0.4809 0.07043 0.06738 -0.0540 0.9922 0.0455 -7.750 -0.4638 0.06570 0.06254 -0.0600 0.9848 0.0474 -7.500 -0.4542 0.06030 0.05667 -0.0664 0.9721 0.0518 -7.250 -0.4454 0.05442 0.05052 -0.0698 0.9622 0.0545 -7.000 -0.4213 0.05181 0.04797 -0.0713 0.9552 0.0579 -6.500 -0.3787 0.04453 0.04021 -0.0759 0.9383 0.0693 -6.250 -0.3493 0.04077 0.03590 -0.0793 0.9332 0.0784 -6.000 -0.3260 0.03792 0.03305 -0.0798 0.9231 0.0816 -5.750 -0.2834 0.02969 0.02307 -0.0785 0.9195 0.0350 -5.500 -0.2559 0.02605 0.01935 -0.0785 0.9115 0.0309 -5.250 -0.2174 0.02341 0.01617 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0.01629 0.00931 -0.0569 0.5395 0.9491 5.000 0.8471 0.01638 0.00934 -0.0567 0.5332 0.9521 5.250 0.8662 0.01625 0.00932 -0.0553 0.5237 0.9556 5.500 0.8869 0.01618 0.00924 -0.0542 0.5139 0.9593 5.750 0.9036 0.01610 0.00914 -0.0523 0.5038 0.9630 6.000 0.9209 0.01598 0.00909 -0.0506 0.4930 0.9663 6.250 0.9445 0.01596 0.00908 -0.0501 0.4838 0.9697 6.500 0.9626 0.01592 0.00907 -0.0486 0.4742 0.9735 6.750 0.9810 0.01588 0.00911 -0.0472 0.4633 0.9768 7.000 1.0053 0.01590 0.00916 -0.0470 0.4532 0.9801 7.250 1.0253 0.01592 0.00921 -0.0460 0.4427 0.9838 7.500 1.0462 0.01597 0.00936 -0.0453 0.4315 0.9875 7.750 1.0673 0.01604 0.00949 -0.0446 0.4198 0.9915 8.000 1.0883 0.01618 0.00965 -0.0439 0.4098 0.9950 8.250 1.1131 0.01633 0.00989 -0.0441 0.3977 0.9978 8.500 1.1256 0.01642 0.01008 -0.0418 0.3871 0.9999 8.750 1.1381 0.01663 0.01034 -0.0397 0.3763 0.9999 9.000 1.1539 0.01692 0.01068 -0.0383 0.3640 0.9999 9.250 1.1708 0.01731 0.01110 -0.0372 0.3503 0.9999 9.500 1.1892 0.01776 0.01164 -0.0365 0.3355 0.9999 9.750 1.2068 0.01830 0.01224 -0.0357 0.3189 0.9999 10.000 1.2242 0.01892 0.01293 -0.0351 0.3004 0.9999 10.250 1.2389 0.01973 0.01375 -0.0342 0.2793 0.9999 10.500 1.2492 0.02083 0.01481 -0.0331 0.2493 0.9999 10.750 1.2543 0.02235 0.01622 -0.0316 0.2168 0.9999 11.000 1.2553 0.02433 0.01804 -0.0302 0.1788 0.9999 11.250 1.2532 0.02675 0.02029 -0.0289 0.1418 0.9999 11.500 1.2493 0.02952 0.02292 -0.0280 0.1103 0.9999 11.750 1.2443 0.03261 0.02589 -0.0275 0.0865 0.9999 12.000 1.2408 0.03574 0.02898 -0.0273 0.0662 0.9999 12.250 1.2357 0.03917 0.03236 -0.0273 0.0571 0.9999 12.500 1.2361 0.04224 0.03554 -0.0278 0.0500 0.9999 12.750 1.2334 0.04570 0.03908 -0.0283 0.0455 0.9999 13.000 1.2344 0.04884 0.04234 -0.0290 0.0424 0.9999 13.250 1.2363 0.05194 0.04557 -0.0297 0.0398 0.9999 13.500 1.2357 0.05535 0.04906 -0.0306 0.0378 0.9999 13.750 1.2312 0.05927 0.05305 -0.0315 0.0361 0.9999 14.000 1.2323 0.06263 0.05653 -0.0325 0.0348 0.9999 14.250 1.2328 0.06606 0.06009 -0.0335 0.0336 0.9999 14.500 1.2327 0.06955 0.06369 -0.0345 0.0326 0.9999 14.750 1.2328 0.07313 0.06738 -0.0356 0.0316 0.9999 15.000 1.2320 0.07676 0.07110 -0.0366 0.0306 0.9999 15.250 1.2307 0.08043 0.07483 -0.0376 0.0296 0.9999 15.500 1.2290 0.08376 0.07817 -0.0380 0.0282 0.9999 15.750 1.2325 0.08719 0.08181 -0.0394 0.0270 0.9999