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 = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -10.250 -0.5076 0.08358 0.08146 -0.0373 1.0000 0.0146 -10.000 -0.5237 0.07721 0.07512 -0.0398 1.0000 0.0144 -9.750 -0.5622 0.06691 0.06478 -0.0459 1.0000 0.0135 -9.500 -0.5958 0.06127 0.05906 -0.0453 1.0000 0.0131 -9.250 -0.6255 0.05767 0.05537 -0.0417 1.0000 0.0129 -9.000 -0.6526 0.05533 0.05295 -0.0363 1.0000 0.0129 -8.750 -0.6795 0.05376 0.05133 -0.0295 1.0000 0.0128 -8.500 -0.7019 0.05164 0.04914 -0.0233 1.0000 0.0128 -8.250 -0.7190 0.04913 0.04655 -0.0181 1.0000 0.0130 -8.000 -0.7320 0.04657 0.04390 -0.0133 1.0000 0.0131 -7.750 -0.7341 0.04376 0.04100 -0.0109 0.9991 0.0135 -7.500 -0.7271 0.04003 0.03711 -0.0100 0.9969 0.0141 -7.250 -0.7169 0.03615 0.03305 -0.0093 0.9949 0.0150 -7.000 -0.7073 0.03252 0.02919 -0.0077 0.9925 0.0161 -5.250 -0.6172 0.02020 0.01390 0.0093 0.9833 0.0136 -5.000 -0.5869 0.01859 0.01209 0.0088 0.9827 0.0143 -4.750 -0.5645 0.01753 0.01093 0.0099 0.9800 0.0152 -4.500 -0.5387 0.01659 0.00991 0.0102 0.9778 0.0163 -4.250 -0.5104 0.01605 0.00931 0.0099 0.9758 0.0183 -4.000 -0.4820 0.01527 0.00844 0.0096 0.9741 0.0193 -3.750 -0.4570 0.01400 0.00706 0.0100 0.9726 0.0215 -3.500 -0.4276 0.01333 0.00633 0.0094 0.9713 0.0253 -3.250 -0.3950 0.01306 0.00603 0.0080 0.9702 0.0306 -3.000 -0.3744 0.01249 0.00540 0.0095 0.9667 0.0414 -2.750 -0.3549 0.01157 0.00497 0.0107 0.9631 0.1570 -2.500 -0.3491 0.00922 0.00445 0.0141 0.9599 0.5946 -2.250 -0.3307 0.00836 0.00451 0.0165 0.9581 0.8182 -2.000 -0.2992 0.00847 0.00474 0.0161 0.9570 0.8757 -1.750 -0.2752 0.00870 0.00498 0.0173 0.9539 0.9023 -1.500 -0.2468 0.00917 0.00546 0.0178 0.9512 0.9213 -1.250 -0.2120 0.00964 0.00587 0.0168 0.9498 0.9328 -1.000 -0.1770 0.01006 0.00625 0.0155 0.9486 0.9428 -0.750 -0.1319 0.01031 0.00647 0.0116 0.9486 0.9443 -0.500 -0.0895 0.01043 0.00657 0.0081 0.9481 0.9454 -0.250 -0.0443 0.01057 0.00668 0.0040 0.9480 0.9465 0.000 0.0000 0.01064 0.00674 0.0000 0.9475 0.9475 0.250 0.0444 0.01057 0.00668 -0.0040 0.9465 0.9480 0.500 0.0895 0.01043 0.00657 -0.0081 0.9454 0.9481 0.750 0.1319 0.01031 0.00647 -0.0116 0.9443 0.9486 1.000 0.1765 0.01007 0.00627 -0.0154 0.9429 0.9486 1.250 0.2119 0.00964 0.00588 -0.0168 0.9329 0.9498 1.500 0.2469 0.00916 0.00544 -0.0178 0.9211 0.9511 1.750 0.2752 0.00870 0.00497 -0.0173 0.9023 0.9539 2.000 0.2992 0.00847 0.00474 -0.0160 0.8757 0.9570 2.250 0.3307 0.00836 0.00452 -0.0165 0.8186 0.9581 2.500 0.3491 0.00922 0.00445 -0.0141 0.5947 0.9599 2.750 0.3548 0.01157 0.00497 -0.0107 0.1568 0.9630 3.000 0.3744 0.01248 0.00539 -0.0095 0.0414 0.9666 3.250 0.3949 0.01306 0.00602 -0.0080 0.0306 0.9702 3.500 0.4276 0.01333 0.00633 -0.0094 0.0254 0.9713 3.750 0.4568 0.01401 0.00707 -0.0100 0.0214 0.9726 4.000 0.4819 0.01527 0.00844 -0.0096 0.0194 0.9741 4.250 0.5104 0.01605 0.00931 -0.0099 0.0183 0.9758 4.500 0.5386 0.01659 0.00991 -0.0102 0.0163 0.9778 4.750 0.5644 0.01752 0.01092 -0.0099 0.0152 0.9800 5.250 0.6172 0.02020 0.01390 -0.0093 0.0136 0.9833 6.250 0.6950 0.04014 0.03569 -0.0027 0.0198 0.9887 6.500 0.7115 0.04278 0.03866 -0.0011 0.0197 0.9915 6.750 0.7444 0.04183 0.03801 -0.0002 0.0166 0.9938 7.000 0.7623 0.04505 0.04149 0.0005 0.0151 0.9956 7.250 0.7779 0.04809 0.04474 0.0010 0.0141 0.9976 7.500 0.7910 0.05099 0.04780 0.0013 0.0134 0.9996 7.750 0.7909 0.05335 0.05026 0.0043 0.0130 1.0000 8.000 0.7816 0.05614 0.05316 0.0088 0.0127 1.0000 8.250 0.7662 0.05947 0.05662 0.0140 0.0125 1.0000 8.500 0.7398 0.06378 0.06106 0.0201 0.0122 1.0000 8.750 0.7080 0.06828 0.06568 0.0265 0.0121 1.0000