XFOIL Version 6.96 Calculated polar for: NACA 0010-35 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -11.000 -0.4819 0.09508 0.09354 -0.0338 1.0000 0.0071 -10.750 -0.4910 0.08994 0.08841 -0.0348 1.0000 0.0071 -6.750 -0.6622 0.02302 0.01847 -0.0142 0.9826 0.0074 -6.500 -0.6516 0.02032 0.01545 -0.0108 0.9783 0.0075 -6.250 -0.6273 0.01824 0.01311 -0.0101 0.9767 0.0075 -6.000 -0.6005 0.01654 0.01121 -0.0100 0.9757 0.0074 -5.750 -0.5755 0.01427 0.00872 -0.0096 0.9749 0.0078 -5.500 -0.5474 0.01347 0.00788 -0.0100 0.9738 0.0088 -5.250 -0.5177 0.01294 0.00729 -0.0108 0.9730 0.0097 -5.000 -0.4882 0.01231 0.00660 -0.0115 0.9722 0.0105 -4.750 -0.4583 0.01171 0.00593 -0.0123 0.9715 0.0113 -4.500 -0.4256 0.01149 0.00571 -0.0138 0.9710 0.0120 -4.250 -0.3994 0.01043 0.00452 -0.0137 0.9702 0.0153 -4.000 -0.3683 0.01012 0.00416 -0.0148 0.9695 0.0177 -3.750 -0.3523 0.00995 0.00398 -0.0123 0.9643 0.0195 -3.500 -0.3256 0.00953 0.00353 -0.0122 0.9623 0.0277 -3.250 -0.2966 0.00925 0.00325 -0.0128 0.9608 0.0390 -3.000 -0.2711 0.00864 0.00296 -0.0127 0.9592 0.1221 -2.750 -0.2526 0.00749 0.00259 -0.0115 0.9573 0.3356 -2.500 -0.2470 0.00643 0.00233 -0.0071 0.9520 0.5469 -2.250 -0.2396 0.00555 0.00212 -0.0027 0.9474 0.7195 -2.000 -0.2195 0.00517 0.00207 -0.0009 0.9452 0.8106 -1.750 -0.1913 0.00507 0.00204 -0.0010 0.9438 0.8430 -1.500 -0.1651 0.00505 0.00208 -0.0006 0.9416 0.8683 -1.250 -0.1431 0.00509 0.00214 0.0008 0.9376 0.8881 -1.000 -0.1151 0.00512 0.00219 0.0008 0.9352 0.8985 -0.750 -0.0859 0.00518 0.00226 0.0006 0.9332 0.9089 -0.500 -0.0580 0.00517 0.00223 0.0006 0.9309 0.9160 -0.250 -0.0278 0.00517 0.00224 0.0000 0.9275 0.9191 0.000 0.0000 0.00518 0.00225 0.0000 0.9227 0.9227 0.250 0.0278 0.00517 0.00224 0.0000 0.9191 0.9275 0.500 0.0580 0.00517 0.00223 -0.0006 0.9160 0.9309 0.750 0.0859 0.00518 0.00226 -0.0006 0.9087 0.9332 1.000 0.1151 0.00512 0.00219 -0.0008 0.8985 0.9352 1.250 0.1431 0.00509 0.00214 -0.0008 0.8882 0.9376 1.500 0.1651 0.00505 0.00208 0.0006 0.8681 0.9416 1.750 0.1913 0.00506 0.00204 0.0010 0.8429 0.9438 2.000 0.2195 0.00517 0.00207 0.0009 0.8106 0.9452 2.250 0.2397 0.00555 0.00212 0.0027 0.7202 0.9474 2.500 0.2468 0.00644 0.00234 0.0072 0.5451 0.9521 2.750 0.2525 0.00749 0.00259 0.0115 0.3347 0.9573 3.000 0.2711 0.00864 0.00296 0.0127 0.1219 0.9592 3.250 0.2966 0.00924 0.00325 0.0127 0.0392 0.9608 3.500 0.3257 0.00953 0.00353 0.0122 0.0278 0.9622 3.750 0.3524 0.00995 0.00397 0.0122 0.0195 0.9643 4.000 0.3731 0.01015 0.00420 0.0136 0.0175 0.9683 4.250 0.3994 0.01043 0.00452 0.0137 0.0153 0.9702 4.500 0.4257 0.01148 0.00570 0.0138 0.0120 0.9710 4.750 0.4584 0.01170 0.00593 0.0123 0.0113 0.9715 5.000 0.4883 0.01230 0.00659 0.0115 0.0105 0.9722 5.250 0.5177 0.01292 0.00727 0.0108 0.0097 0.9730 5.500 0.5475 0.01346 0.00786 0.0100 0.0088 0.9738 5.750 0.5754 0.01427 0.00872 0.0096 0.0078 0.9748 6.000 0.6004 0.01662 0.01130 0.0100 0.0074 0.9757 6.250 0.6274 0.01828 0.01315 0.0101 0.0075 0.9768 6.500 0.6517 0.02033 0.01546 0.0107 0.0075 0.9783 6.750 0.6627 0.02297 0.01841 0.0141 0.0074 0.9826 8.000 0.7092 0.04825 0.04572 0.0250 0.0071 0.9921 8.250 0.7108 0.05311 0.05080 0.0272 0.0071 0.9947 8.500 0.7137 0.05847 0.05638 0.0278 0.0071 0.9963 8.750 0.7144 0.06384 0.06193 0.0279 0.0071 0.9984 9.000 0.7043 0.06793 0.06613 0.0299 0.0070 1.0000 9.250 0.6776 0.06936 0.06760 0.0365 0.0071 1.0000 9.500 0.6533 0.07166 0.06994 0.0410 0.0071 1.0000 9.750 0.6314 0.07483 0.07316 0.0435 0.0071 1.0000 10.000 0.6122 0.07908 0.07745 0.0435 0.0071 1.0000