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 = 0.200 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -10.250 -0.3845 0.09142 0.08801 -0.0417 1.0000 0.0111 -10.000 -0.3887 0.08668 0.08331 -0.0439 1.0000 0.0109 -9.750 -0.3959 0.08121 0.07791 -0.0467 1.0000 0.0107 -9.500 -0.4091 0.07396 0.07072 -0.0515 1.0000 0.0105 -9.250 -0.4348 0.06547 0.06221 -0.0576 1.0000 0.0102 -9.000 -0.4832 0.05564 0.05215 -0.0595 0.9822 0.0096 -8.750 -0.4769 0.05015 0.04637 -0.0641 0.9290 0.0095 -8.500 -0.4560 0.04381 0.03951 -0.0703 0.8948 0.0093 -8.250 -0.4389 0.03831 0.03343 -0.0733 0.8600 0.0091 -7.750 -0.4181 0.03010 0.02398 -0.0713 0.8068 0.0086 -7.500 -0.4051 0.02692 0.02019 -0.0693 0.7871 0.0084 -7.250 -0.3879 0.02439 0.01713 -0.0677 0.7701 0.0083 -7.000 -0.3679 0.02257 0.01494 -0.0664 0.7548 0.0082 -6.750 -0.3468 0.02120 0.01329 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0.0352 1.0000 13.250 1.1535 0.05144 0.04447 -0.0044 0.0326 1.0000 13.500 1.1539 0.05434 0.04741 -0.0048 0.0304 1.0000 13.750 1.1547 0.05729 0.05047 -0.0053 0.0290 1.0000 14.000 1.1549 0.06039 0.05368 -0.0059 0.0278 1.0000 14.250 1.1550 0.06358 0.05698 -0.0067 0.0268 1.0000 14.500 1.1545 0.06691 0.06041 -0.0075 0.0258 1.0000 14.750 1.1535 0.07037 0.06396 -0.0084 0.0248 1.0000 15.000 1.1513 0.07406 0.06772 -0.0095 0.0237 1.0000 15.250 1.1489 0.07786 0.07160 -0.0106 0.0225 1.0000 15.500 1.1509 0.08110 0.07497 -0.0117 0.0218 1.0000 15.750 1.1534 0.08434 0.07834 -0.0128 0.0208 1.0000 16.000 1.1551 0.08775 0.08187 -0.0140 0.0198 1.0000 16.250 1.1563 0.09127 0.08550 -0.0153 0.0190 1.0000 16.500 1.1567 0.09493 0.08928 -0.0168 0.0183 1.0000 16.750 1.1565 0.09876 0.09320 -0.0183 0.0176 1.0000 17.000 1.1549 0.10283 0.09735 -0.0200 0.0170 1.0000