XFOIL Version 6.96 Calculated polar for: NASA SC(2)-1010 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -17.000 -0.8409 0.08011 0.07703 -0.0528 0.9999 0.0145 -16.750 -0.8881 0.06657 0.06327 -0.0616 0.9999 0.0144 -16.500 -0.9158 0.05758 0.05409 -0.0679 0.9999 0.0144 -16.250 -0.9350 0.05085 0.04720 -0.0726 0.9999 0.0144 -16.000 -0.9499 0.04559 0.04178 -0.0758 0.9999 0.0144 -15.750 -0.9634 0.04152 0.03757 -0.0773 0.9999 0.0145 -15.500 -0.9789 0.03840 0.03433 -0.0768 0.9999 0.0145 -15.250 -0.9993 0.03607 0.03192 -0.0741 0.9999 0.0146 -15.000 -1.0052 0.03360 0.02930 -0.0740 0.9999 0.0147 -14.750 -0.9817 0.03069 0.02616 -0.0796 0.9977 0.0149 -14.500 -0.9543 0.02840 0.02364 -0.0843 0.9958 0.0151 -14.250 -0.9260 0.02654 0.02158 -0.0880 0.9945 0.0153 -14.000 -0.8984 0.02479 0.01970 -0.0911 0.9930 0.0157 -13.750 -0.8703 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0.8793 0.01332 0.00528 -0.1412 0.0517 0.7135 1.000 0.9056 0.01364 0.00558 -0.1408 0.0459 0.7199 1.250 0.9319 0.01393 0.00588 -0.1404 0.0419 0.7260 1.500 0.9577 0.01427 0.00622 -0.1399 0.0384 0.7329 1.750 0.9839 0.01456 0.00654 -0.1394 0.0361 0.7397 2.000 1.0095 0.01489 0.00690 -0.1389 0.0336 0.7457 2.250 1.0349 0.01528 0.00729 -0.1383 0.0315 0.7517 2.500 1.0606 0.01560 0.00764 -0.1378 0.0297 0.7568 2.750 1.0858 0.01596 0.00804 -0.1371 0.0281 0.7617 3.000 1.1104 0.01640 0.00848 -0.1364 0.0266 0.7673 3.250 1.1351 0.01682 0.00895 -0.1356 0.0255 0.7728 3.500 1.1595 0.01722 0.00940 -0.1349 0.0243 0.7784 3.750 1.1840 0.01763 0.00986 -0.1341 0.0232 0.7849 4.000 1.2077 0.01809 0.01033 -0.1332 0.0221 0.7904 4.250 1.2308 0.01863 0.01091 -0.1322 0.0211 0.7959 4.500 1.2544 0.01911 0.01146 -0.1313 0.0204 0.8011 4.750 1.2774 0.01962 0.01204 -0.1303 0.0197 0.8053 5.000 1.3001 0.02015 0.01262 -0.1292 0.0190 0.8099 5.250 1.3226 0.02069 0.01321 -0.1282 0.0184 0.8148 5.500 1.3444 0.02131 0.01387 -0.1270 0.0179 0.8192 5.750 1.3643 0.02211 0.01472 -0.1255 0.0174 0.8240 6.000 1.3854 0.02277 0.01548 -0.1242 0.0170 0.8296 6.250 1.4059 0.02347 0.01627 -0.1228 0.0167 0.8345 6.500 1.4259 0.02417 0.01708 -0.1213 0.0162 0.8391 6.750 1.4457 0.02487 0.01787 -0.1198 0.0158 0.8440 7.000 1.4653 0.02556 0.01863 -0.1183 0.0154 0.8487 7.250 1.4838 0.02627 0.01943 -0.1167 0.0150 0.8531 7.500 1.5015 0.02703 0.02026 -0.1149 0.0147 0.8579 7.750 1.5177 0.02789 0.02121 -0.1129 0.0145 0.8628 8.000 1.5315 0.02886 0.02228 -0.1106 0.0142 0.8679 8.250 1.5424 0.03010 0.02361 -0.1078 0.0139 0.8738 8.500 1.5539 0.03109 0.02476 -0.1050 0.0138 0.8799 8.750 1.5632 0.03216 0.02599 -0.1020 0.0137 0.8872 9.000 1.5723 0.03334 0.02735 -0.0991 0.0135 0.8950 9.250 1.5804 0.03461 0.02879 -0.0961 0.0133 0.9039 9.500 1.5868 0.03596 0.03032 -0.0930 0.0131 0.9154 10.000 1.5939 0.03893 0.03364 -0.0864 0.0128 1.0001 10.250 1.5974 0.04074 0.03561 -0.0836 0.0127 1.0001 10.500 1.5997 0.04270 0.03773 -0.0809 0.0126 1.0001 10.750 1.6004 0.04482 0.04001 -0.0783 0.0124 1.0001 11.000 1.6001 0.04708 0.04243 -0.0758 0.0123 1.0001 11.250 1.5986 0.04952 0.04502 -0.0736 0.0122 1.0001 11.500 1.5952 0.05223 0.04789 -0.0716 0.0121 1.0001 11.750 1.5914 0.05510 0.05091 -0.0699 0.0119 1.0001 12.000 1.5850 0.05841 0.05437 -0.0686 0.0118 1.0001 12.250 1.5764 0.06216 0.05828 -0.0677 0.0118 1.0001 12.500 1.5665 0.06633 0.06261 -0.0674 0.0117 1.0001 12.750 1.5551 0.07099 0.06743 -0.0678 0.0116 1.0001 13.000 1.5386 0.07687 0.07349 -0.0692 0.0116 1.0001 13.250 1.5201 0.08379 0.08061 -0.0720 0.0115 1.0001 13.500 1.4919 0.09352 0.09059 -0.0773 0.0115 1.0001 13.750 1.4522 0.10717 0.10454 -0.0861 0.0116 1.0001