XFOIL Version 6.96 Calculated polar for: FX 61-147 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 ------ -------- --------- --------- -------- -------- -------- -9.500 -0.3865 0.08999 0.08681 -0.0519 1.0000 0.0495 -9.250 -0.4112 0.08424 0.08110 -0.0552 1.0000 0.0497 -9.000 -0.4309 0.08055 0.07744 -0.0551 1.0000 0.0496 -8.750 -0.4562 0.07786 0.07478 -0.0536 1.0000 0.0495 -8.500 -0.4739 0.07547 0.07245 -0.0511 1.0000 0.0492 -8.250 -0.5085 0.07382 0.07081 -0.0486 0.9994 0.0493 -8.000 -0.5006 0.06800 0.06434 -0.0632 0.9867 0.0503 -7.750 -0.4924 0.05870 0.05513 -0.0668 0.9812 0.0522 -7.500 -0.4628 0.05593 0.05247 -0.0694 0.9771 0.0559 -7.250 -0.3783 0.03425 0.03055 -0.0810 0.9526 0.0669 -7.000 -0.3543 0.03047 0.02663 -0.0838 0.9466 0.0709 -6.750 -0.3349 0.02603 0.02174 -0.0872 0.9384 0.0801 -6.500 -0.3525 0.03281 0.02707 -0.0846 0.9499 0.0369 -6.250 -0.3245 0.02798 0.02166 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3.750 0.8917 0.01651 0.00910 -0.1029 0.5341 0.8146 4.000 0.9197 0.01663 0.00929 -0.1034 0.5264 0.8182 4.250 0.9453 0.01673 0.00939 -0.1030 0.5201 0.8212 4.500 0.9700 0.01687 0.00961 -0.1026 0.5132 0.8247 4.750 0.9982 0.01702 0.00977 -0.1030 0.5064 0.8283 5.000 1.0277 0.01724 0.01004 -0.1038 0.4996 0.8318 5.250 1.0521 0.01735 0.01021 -0.1033 0.4929 0.8348 5.500 1.0786 0.01755 0.01043 -0.1033 0.4870 0.8381 5.750 1.1037 0.01771 0.01070 -0.1031 0.4795 0.8415 6.000 1.1344 0.01797 0.01093 -0.1040 0.4734 0.8450 6.250 1.1576 0.01813 0.01126 -0.1034 0.4659 0.8482 6.500 1.1841 0.01830 0.01143 -0.1033 0.4595 0.8514 6.750 1.2064 0.01845 0.01173 -0.1025 0.4508 0.8549 7.000 1.2340 0.01864 0.01193 -0.1028 0.4430 0.8585 7.250 1.2565 0.01876 0.01220 -0.1021 0.4337 0.8620 7.500 1.2779 0.01887 0.01238 -0.1010 0.4247 0.8653 7.750 1.3009 0.01898 0.01252 -0.1003 0.4154 0.8688 8.000 1.3217 0.01913 0.01284 -0.0993 0.4050 0.8729 8.250 1.3416 0.01923 0.01303 -0.0981 0.3934 0.8766 8.500 1.3583 0.01930 0.01317 -0.0962 0.3811 0.8804 8.750 1.3757 0.01944 0.01337 -0.0946 0.3683 0.8845 9.000 1.3942 0.01969 0.01369 -0.0933 0.3559 0.8890 9.250 1.4084 0.01990 0.01405 -0.0911 0.3434 0.8936 9.500 1.4217 0.02017 0.01446 -0.0888 0.3299 0.8986 9.750 1.4332 0.02055 0.01489 -0.0863 0.3140 0.9038 10.000 1.4388 0.02098 0.01537 -0.0828 0.2963 0.9103 10.250 1.4445 0.02161 0.01603 -0.0796 0.2736 0.9181 10.500 1.4443 0.02237 0.01678 -0.0755 0.2501 0.9281 10.750 1.4399 0.02338 0.01776 -0.0711 0.2240 0.9447 11.000 1.4362 0.02485 0.01914 -0.0678 0.1920 1.0000 11.250 1.4297 0.02723 0.02132 -0.0653 0.1558 1.0000 11.500 1.4208 0.03004 0.02396 -0.0630 0.1257 1.0000 11.750 1.4109 0.03319 0.02698 -0.0613 0.1015 1.0000 12.000 1.4021 0.03651 0.03023 -0.0601 0.0830 1.0000 12.250 1.3938 0.04003 0.03374 -0.0595 0.0697 1.0000 12.500 1.3865 0.04366 0.03739 -0.0592 0.0603 1.0000 12.750 1.3765 0.04778 0.04150 -0.0592 0.0537 1.0000 13.000 1.3730 0.05138 0.04521 -0.0596 0.0472 1.0000 13.250 1.3629 0.05586 0.04972 -0.0601 0.0424 1.0000 13.500 1.3584 0.05982 0.05378 -0.0608 0.0369 1.0000 13.750 1.3455 0.06499 0.05900 -0.0619 0.0329 1.0000 14.000 1.3373 0.06972 0.06384 -0.0630 0.0281 1.0000 14.250 1.3191 0.07590 0.07006 -0.0645 0.0255 1.0000 14.500 1.3138 0.08052 0.07481 -0.0657 0.0224 1.0000 14.750 1.3066 0.08544 0.07980 -0.0672 0.0205 1.0000 15.000 1.2965 0.09071 0.08510 -0.0686 0.0192 1.0000