XFOIL Version 6.96 Calculated polar for: FX 63-143 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -10.250 -0.3677 0.10202 0.09526 -0.0432 1.0000 0.0397 -10.000 -0.3780 0.09547 0.08880 -0.0472 1.0000 0.0372 -9.750 -0.3792 0.09131 0.08473 -0.0488 1.0000 0.0367 -9.500 -0.3837 0.08648 0.07998 -0.0512 1.0000 0.0361 -9.250 -0.3938 0.08126 0.07483 -0.0541 1.0000 0.0355 -9.000 -0.4102 0.07662 0.07025 -0.0558 1.0000 0.0346 -8.750 -0.4455 0.07196 0.06556 -0.0554 1.0000 0.0332 -8.500 -0.4634 0.06940 0.06305 -0.0530 1.0000 0.0331 -8.250 -0.4767 0.06694 0.06062 -0.0505 1.0000 0.0328 -8.000 -0.4936 0.06467 0.05836 -0.0473 1.0000 0.0327 -7.750 -0.5114 0.06317 0.05689 -0.0432 1.0000 0.0324 -7.500 -0.4989 0.05834 0.05170 -0.0464 0.9801 0.0323 -7.250 -0.4841 0.05387 0.04680 -0.0485 0.9583 0.0322 -7.000 -0.4646 0.04972 0.04216 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1.0000 2.750 0.6574 0.02819 0.01658 -0.0647 0.6036 1.0000 3.000 0.6698 0.02872 0.01713 -0.0626 0.5965 1.0000 3.250 0.6920 0.02886 0.01722 -0.0615 0.5917 1.0000 3.500 0.6949 0.02982 0.01827 -0.0582 0.5837 1.0000 3.750 0.7092 0.03035 0.01883 -0.0563 0.5784 1.0000 4.000 0.7310 0.03062 0.01912 -0.0553 0.5747 1.0000 4.250 0.7252 0.03198 0.02060 -0.0510 0.5668 1.0000 4.500 0.7388 0.03256 0.02122 -0.0490 0.5618 1.0000 4.750 0.7629 0.03273 0.02140 -0.0482 0.5582 1.0000 5.000 0.7483 0.03434 0.02312 -0.0428 0.5490 1.0000 5.250 0.7700 0.03448 0.02331 -0.0415 0.5438 1.0000 5.500 0.7655 0.03565 0.02455 -0.0373 0.5359 1.0000 5.750 0.7723 0.03639 0.02535 -0.0344 0.5292 1.0000 6.000 0.8018 0.03629 0.02532 -0.0341 0.5252 1.0000 6.250 0.7713 0.03865 0.02775 -0.0274 0.5146 1.0000 6.500 0.7963 0.03876 0.02794 -0.0267 0.5100 1.0000 6.750 0.7728 0.04120 0.03043 -0.0217 0.4998 1.0000 7.000 0.7942 0.04155 0.03088 -0.0208 0.4944 1.0000 7.250 0.7754 0.04404 0.03341 -0.0169 0.4844 1.0000 7.500 0.7982 0.04417 0.03365 -0.0160 0.4778 1.0000 7.750 0.7863 0.04674 0.03628 -0.0136 0.4671 1.0000 8.000 0.8175 0.04596 0.03563 -0.0128 0.4601 1.0000 8.250 0.8039 0.04902 0.03875 -0.0112 0.4477 1.0000 8.500 0.8529 0.04628 0.03617 -0.0103 0.4418 1.0000 8.750 0.8354 0.04984 0.03980 -0.0090 0.4286 1.0000 9.000 0.8931 0.04612 0.03629 -0.0080 0.4240 1.0000 9.250 0.8688 0.05052 0.04076 -0.0070 0.4101 1.0000 9.750 0.8885 0.05306 0.04359 -0.0056 0.3919 1.0000 10.250 0.9081 0.05578 0.04662 -0.0044 0.3743 1.0000 10.750 0.9005 0.06178 0.05287 -0.0042 0.3505 1.0000 11.500 0.9343 0.06527 0.05690 -0.0028 0.3214 1.0000 11.750 0.9428 0.06676 0.05858 -0.0025 0.3105 1.0000 12.250 0.9556 0.07027 0.06245 -0.0020 0.2864 1.0000 12.500 0.9752 0.06998 0.06238 -0.0011 0.2740 1.0000 12.750 0.9730 0.07300 0.06555 -0.0014 0.2585 1.0000 13.000 0.9807 0.07445 0.06717 -0.0011 0.2420 1.0000 13.250 1.0101 0.07198 0.06467 0.0009 0.2171 1.0000 13.500 0.9975 0.07665 0.06943 -0.0002 0.1948 1.0000 13.750 1.0048 0.07766 0.07009 0.0003 0.1681 1.0000 14.000 0.9973 0.08182 0.07424 -0.0007 0.1474 1.0000 14.250 0.9957 0.08488 0.07704 -0.0013 0.1325 1.0000 14.500 0.9923 0.08854 0.08058 -0.0021 0.1193 1.0000 14.750 0.9875 0.09281 0.08493 -0.0035 0.1062 1.0000 15.000 0.9858 0.09645 0.08848 -0.0046 0.0975 1.0000 15.250 0.9869 0.09962 0.09155 -0.0053 0.0903 1.0000