XFOIL Version 6.96 Calculated polar for: NACA 63A210 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.750 -0.5685 0.08995 0.08604 -0.0045 0.6685 0.0343 -8.500 -0.5703 0.08432 0.08043 -0.0096 0.6691 0.0349 -8.250 -0.5759 0.07758 0.07369 -0.0175 0.6697 0.0349 -8.000 -0.5825 0.07259 0.06861 -0.0214 0.6703 0.0353 -7.750 -0.5834 0.06781 0.06369 -0.0243 0.6708 0.0362 -7.500 -0.5800 0.06354 0.05918 -0.0265 0.6713 0.0380 -7.250 -0.5714 0.06246 0.05754 -0.0270 0.6719 0.0398 -7.000 -0.5671 0.05730 0.05205 -0.0274 0.6723 0.0404 -6.750 -0.5583 0.05041 0.04524 -0.0283 0.6727 0.0423 -6.500 -0.5429 0.04728 0.04205 -0.0284 0.6732 0.0444 -6.250 -0.5261 0.04442 0.03898 -0.0283 0.6736 0.0478 -6.000 -0.5101 0.04219 0.03595 -0.0273 0.6739 0.0550 -4.500 -0.3619 0.02484 0.01673 -0.0226 0.6762 0.0384 -4.250 -0.3345 0.02198 0.01362 -0.0219 0.6766 0.0340 -4.000 -0.3065 0.02013 0.01156 -0.0212 0.6769 0.0317 -3.750 -0.2798 0.01873 0.01005 -0.0205 0.6773 0.0314 -3.500 -0.2543 0.01758 0.00888 -0.0199 0.6777 0.0327 -3.250 -0.2292 0.01668 0.00795 -0.0193 0.6777 0.0358 -3.000 -0.2044 0.01588 0.00709 -0.0189 0.6778 0.0423 -2.750 -0.1789 0.01515 0.00625 -0.0184 0.6780 0.0532 -2.500 -0.1734 0.01200 0.00564 -0.0151 0.6783 0.6246 -2.250 -0.1539 0.01191 0.00588 -0.0125 0.6785 0.7405 -2.000 -0.1356 0.01197 0.00615 -0.0091 0.6788 0.8097 -1.750 -0.1136 0.01201 0.00622 -0.0071 0.6791 0.8448 -1.500 -0.0870 0.01201 0.00616 -0.0067 0.6793 0.8608 -1.250 -0.0600 0.01202 0.00613 -0.0065 0.6797 0.8766 -1.000 -0.0323 0.01206 0.00611 -0.0063 0.6800 0.8924 -0.750 -0.0032 0.01213 0.00615 -0.0065 0.6804 0.9076 -0.500 0.0281 0.01223 0.00623 -0.0073 0.6808 0.9226 -0.250 0.0619 0.01236 0.00632 -0.0086 0.6804 0.9364 0.000 0.0997 0.01252 0.00642 -0.0108 0.6791 0.9477 0.250 0.1393 0.01269 0.00654 -0.0135 0.6784 0.9583 0.500 0.1801 0.01285 0.00669 -0.0166 0.6783 0.9688 0.750 0.2238 0.01301 0.00686 -0.0204 0.6783 0.9776 1.000 0.2670 0.01320 0.00703 -0.0242 0.6769 0.9860 1.250 0.3096 0.01344 0.00726 -0.0278 0.6750 0.9948 1.500 0.3466 0.01367 0.00751 -0.0306 0.6738 1.0000 1.750 0.3690 0.01387 0.00778 -0.0305 0.6733 1.0000 2.000 0.3918 0.01411 0.00808 -0.0304 0.6724 1.0000 2.250 0.4149 0.01437 0.00840 -0.0303 0.6694 1.0000 2.500 0.4377 0.01467 0.00872 -0.0297 0.6640 1.0000 2.750 0.4602 0.01512 0.00915 -0.0288 0.6577 1.0000 3.000 0.4824 0.01545 0.00952 -0.0277 0.6476 1.0000 3.250 0.5042 0.01585 0.00994 -0.0266 0.6359 1.0000 3.500 0.5256 0.01633 0.01042 -0.0253 0.6227 1.0000 3.750 0.5470 0.01678 0.01088 -0.0239 0.6089 1.0000 4.000 0.5688 0.01717 0.01133 -0.0226 0.5950 1.0000 4.250 0.5911 0.01753 0.01171 -0.0213 0.5814 1.0000 4.500 0.6135 0.01796 0.01210 -0.0198 0.5672 1.0000 4.750 0.6349 0.01863 0.01282 -0.0185 0.5512 1.0000 5.000 0.6524 0.02035 0.01457 -0.0168 0.5302 1.0000 5.250 0.6672 0.02248 0.01675 -0.0150 0.5019 1.0000 5.500 0.6888 0.02259 0.01708 -0.0138 0.4842 1.0000 5.750 0.7313 0.01730 0.01195 -0.0140 0.4402 1.0000 6.000 0.7538 0.01679 0.01057 -0.0131 0.2400 1.0000 6.250 0.7724 0.01750 0.01067 -0.0120 0.1483 1.0000 6.500 0.7922 0.01832 0.01127 -0.0112 0.1088 1.0000 6.750 0.8117 0.01924 0.01208 -0.0102 0.0773 1.0000 7.000 0.8292 0.02041 0.01311 -0.0088 0.0503 1.0000 7.250 0.8469 0.02153 0.01419 -0.0075 0.0369 1.0000 7.500 0.8603 0.02310 0.01579 -0.0053 0.0310 1.0000 7.750 0.8786 0.02416 0.01691 -0.0040 0.0261 1.0000 8.000 0.8884 0.02632 0.01908 -0.0015 0.0226 1.0000 8.250 0.9061 0.02783 0.02077 0.0002 0.0214 1.0000 8.500 0.9232 0.02983 0.02297 0.0020 0.0204 1.0000 8.750 0.9406 0.03233 0.02573 0.0036 0.0199 1.0000 9.000 0.9567 0.03545 0.02919 0.0052 0.0198 1.0000 9.250 0.9685 0.03922 0.03337 0.0069 0.0201 1.0000 9.500 0.9744 0.04349 0.03811 0.0086 0.0207 1.0000 9.750 0.9740 0.04813 0.04313 0.0101 0.0214 1.0000 10.000 0.9666 0.05334 0.04865 0.0114 0.0222 1.0000