XFOIL Version 6.96 Calculated polar for: NASA SC(2)-0606 AIRFOIL 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 ------ -------- --------- --------- -------- -------- -------- -11.250 -0.5077 0.10055 0.09893 -0.0118 1.0000 0.0099 -11.000 -0.5077 0.09669 0.09507 -0.0127 1.0000 0.0101 -10.750 -0.6146 0.10655 0.10484 -0.0020 1.0000 0.0093 -10.500 -0.6132 0.10197 0.10027 -0.0041 1.0000 0.0093 -10.250 -0.6122 0.09728 0.09559 -0.0063 1.0000 0.0094 -10.000 -0.6275 0.08892 0.08728 -0.0105 1.0000 0.0096 -9.750 -0.6328 0.08337 0.08174 -0.0137 1.0000 0.0098 -9.500 -0.6373 0.07817 0.07657 -0.0169 1.0000 0.0099 -9.250 -0.6465 0.07251 0.07095 -0.0206 1.0000 0.0099 -9.000 -0.6510 0.06240 0.06079 -0.0334 1.0000 0.0098 -8.750 -0.6532 0.05559 0.05385 -0.0381 1.0000 0.0099 -8.500 -0.6491 0.05018 0.04829 -0.0410 1.0000 0.0101 -8.250 -0.6388 0.04607 0.04403 -0.0429 1.0000 0.0103 -8.000 -0.6249 0.04172 0.03949 -0.0447 1.0000 0.0106 -7.750 -0.6047 0.03030 0.02744 -0.0483 1.0000 0.0091 -7.500 -0.5755 0.02292 0.01928 -0.0521 1.0000 0.0095 -7.250 -0.5447 0.01862 0.01446 -0.0546 1.0000 0.0096 -7.000 -0.5144 0.01577 0.01122 -0.0561 1.0000 0.0097 -6.750 -0.4840 0.01348 0.00865 -0.0573 1.0000 0.0100 -6.500 -0.4552 0.01234 0.00741 -0.0580 1.0000 0.0107 -6.250 -0.4262 0.01142 0.00639 -0.0586 1.0000 0.0111 -6.000 -0.3975 0.01075 0.00565 -0.0590 1.0000 0.0119 -5.750 -0.3688 0.01019 0.00502 -0.0594 1.0000 0.0127 -5.500 -0.3398 0.00969 0.00446 -0.0599 1.0000 0.0134 -5.250 -0.3110 0.00928 0.00399 -0.0603 1.0000 0.0140 -5.000 -0.2823 0.00893 0.00360 -0.0607 1.0000 0.0145 -4.750 -0.2502 0.00828 0.00286 -0.0618 1.0000 0.0182 -4.500 -0.2220 0.00809 0.00264 -0.0620 1.0000 0.0214 -4.250 -0.1921 0.00778 0.00235 -0.0626 1.0000 0.0331 -4.000 -0.1629 0.00757 0.00219 -0.0631 1.0000 0.0484 -3.750 -0.1334 0.00737 0.00208 -0.0637 1.0000 0.0709 -3.500 -0.1002 0.00693 0.00193 -0.0653 1.0000 0.1499 -3.250 -0.0604 0.00612 0.00176 -0.0687 1.0000 0.3339 -3.000 -0.0222 0.00548 0.00167 -0.0716 1.0000 0.5034 -2.750 0.0090 0.00532 0.00168 -0.0725 1.0000 0.5697 -2.500 0.0386 0.00525 0.00171 -0.0730 1.0000 0.6051 -2.250 0.0674 0.00523 0.00175 -0.0733 1.0000 0.6288 -2.000 0.0961 0.00522 0.00181 -0.0735 1.0000 0.6526 -1.750 0.1246 0.00522 0.00189 -0.0736 1.0000 0.6771 -1.500 0.1527 0.00525 0.00198 -0.0737 1.0000 0.6954 -1.250 0.1895 0.00515 0.00192 -0.0757 0.9975 0.7097 -1.000 0.2388 0.00481 0.00165 -0.0804 0.9892 0.7253 -0.750 0.2849 0.00454 0.00145 -0.0843 0.9795 0.7368 -0.500 0.3209 0.00442 0.00138 -0.0859 0.9655 0.7457 -0.250 0.3579 0.00436 0.00131 -0.0877 0.9378 0.7544 0.000 0.3812 0.00478 0.00131 -0.0861 0.8053 0.7617 0.250 0.4007 0.00613 0.00154 -0.0844 0.5250 0.7690 0.500 0.4239 0.00732 0.00179 -0.0840 0.2688 0.7753 1.000 0.4767 0.00844 0.00219 -0.0837 0.0700 0.7879 1.250 0.5042 0.00872 0.00238 -0.0837 0.0456 0.7943 1.500 0.5316 0.00897 0.00259 -0.0835 0.0306 0.8002 1.750 0.5590 0.00933 0.00295 -0.0833 0.0196 0.8066 2.000 0.5864 0.00961 0.00328 -0.0831 0.0174 0.8126 2.250 0.6137 0.00990 0.00364 -0.0829 0.0154 0.8188 2.500 0.6406 0.01037 0.00416 -0.0826 0.0136 0.8252 2.750 0.6655 0.01142 0.00537 -0.0818 0.0123 0.8311 3.000 0.6923 0.01194 0.00596 -0.0815 0.0120 0.8377 3.250 0.7183 0.01255 0.00666 -0.0810 0.0117 0.8433 3.500 0.7445 0.01325 0.00747 -0.0805 0.0113 0.8497 3.750 0.7704 0.01406 0.00839 -0.0799 0.0107 0.8554 4.000 0.7962 0.01499 0.00944 -0.0793 0.0101 0.8617 4.250 0.8214 0.01638 0.01102 -0.0785 0.0097 0.8677 4.500 0.8462 0.01779 0.01264 -0.0777 0.0094 0.8736 4.750 0.8715 0.01852 0.01349 -0.0773 0.0089 0.8798 5.000 0.8946 0.02036 0.01560 -0.0763 0.0087 0.8857 5.750 0.9220 0.04687 0.04441 -0.0651 0.0096 0.9042 6.000 0.9342 0.05102 0.04878 -0.0638 0.0092 0.9115 9.250 0.7425 0.11824 0.11706 -0.0708 0.0089 1.0000 9.500 0.7392 0.12273 0.12155 -0.0722 0.0088 1.0000