XFOIL Version 6.96 Calculated polar for: EPPLER 428 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.500 -0.5210 0.10005 0.09223 0.0243 1.0000 0.4121 -8.250 -0.4946 0.09577 0.08786 0.0246 1.0000 0.4258 -8.000 -0.6198 0.07761 0.07009 0.0014 1.0000 0.2773 -7.750 -0.6201 0.07189 0.06435 -0.0013 1.0000 0.2725 -7.500 -0.6341 0.06471 0.05715 -0.0059 1.0000 0.2693 -7.250 -0.6493 0.05716 0.04946 -0.0107 1.0000 0.2690 -7.000 -0.6562 0.05050 0.04250 -0.0141 1.0000 0.2723 -6.750 -0.6514 0.04561 0.03728 -0.0156 1.0000 0.2786 -6.500 -0.6336 0.04349 0.03505 -0.0150 1.0000 0.2889 -6.250 -0.6174 0.04092 0.03231 -0.0147 1.0000 0.2988 -6.000 -0.6035 0.03781 0.02878 -0.0152 1.0000 0.3107 -5.750 -0.5813 0.03664 0.02765 -0.0139 1.0000 0.3227 -5.500 -0.5609 0.03517 0.02613 -0.0130 1.0000 0.3356 -5.250 -0.5410 0.03350 0.02431 -0.0123 1.0000 0.3490 -5.000 -0.5209 0.03171 0.02230 -0.0120 1.0000 0.3624 -4.750 -0.4988 0.03047 0.02101 -0.0111 1.0000 0.3748 -4.500 -0.4769 0.02926 0.01980 -0.0102 1.0000 0.3870 -4.250 -0.4550 0.02790 0.01830 -0.0096 1.0000 0.3995 -4.000 -0.4329 0.02670 0.01698 -0.0090 1.0000 0.4134 -3.750 -0.4107 0.02572 0.01594 -0.0082 1.0000 0.4280 -3.500 -0.3886 0.02486 0.01514 -0.0070 1.0000 0.4416 -3.250 -0.3667 0.02403 0.01435 -0.0060 1.0000 0.4567 -3.000 -0.3451 0.02327 0.01362 -0.0050 1.0000 0.4729 -2.750 -0.3240 0.02259 0.01298 -0.0038 1.0000 0.4904 -2.500 -0.3033 0.02196 0.01241 -0.0027 1.0000 0.5093 -2.250 -0.2828 0.02138 0.01187 -0.0016 1.0000 0.5304 -2.000 -0.2640 0.02089 0.01154 -0.0001 1.0000 0.5518 -1.750 -0.2464 0.02047 0.01125 0.0016 1.0000 0.5755 -1.500 -0.2294 0.02010 0.01101 0.0033 1.0000 0.6039 -1.250 -0.2131 0.01979 0.01086 0.0051 1.0000 0.6370 -1.000 -0.1988 0.01953 0.01081 0.0074 1.0000 0.6761 -0.750 -0.1867 0.01928 0.01083 0.0103 1.0000 0.7268 -0.500 -0.1781 0.01900 0.01087 0.0141 1.0000 0.7922 -0.250 -0.1553 0.01873 0.01098 0.0161 1.0000 0.8964 0.000 -0.0713 0.01866 0.01077 0.0021 1.0000 1.0000 0.250 -0.0368 0.01910 0.01098 -0.0029 0.9887 1.0000 0.500 0.0633 0.01978 0.01155 -0.0172 0.9528 1.0000 0.750 0.1596 0.01981 0.01161 -0.0292 0.9156 1.0000 1.000 0.2449 0.01938 0.01123 -0.0377 0.8757 1.0000 1.250 0.2912 0.01915 0.01094 -0.0390 0.8319 1.0000 1.500 0.3190 0.01911 0.01077 -0.0368 0.7875 1.0000 1.750 0.3375 0.01926 0.01074 -0.0332 0.7398 1.0000 2.000 0.3555 0.01948 0.01074 -0.0295 0.6900 1.0000 2.250 0.3743 0.01980 0.01078 -0.0262 0.6386 1.0000 2.500 0.3946 0.02024 0.01093 -0.0235 0.5900 1.0000 2.750 0.4165 0.02078 0.01117 -0.0214 0.5484 1.0000 3.000 0.4394 0.02140 0.01155 -0.0199 0.5139 1.0000 3.250 0.4631 0.02213 0.01215 -0.0187 0.4852 1.0000 3.500 0.4875 0.02286 0.01268 -0.0177 0.4621 1.0000 3.750 0.5121 0.02367 0.01335 -0.0168 0.4420 1.0000 4.000 0.5369 0.02459 0.01423 -0.0161 0.4247 1.0000 4.250 0.5619 0.02551 0.01507 -0.0154 0.4094 1.0000 4.500 0.5862 0.02654 0.01616 -0.0148 0.3945 1.0000 4.750 0.6102 0.02777 0.01753 -0.0144 0.3818 1.0000 5.000 0.6341 0.02904 0.01886 -0.0138 0.3704 1.0000 5.250 0.6594 0.03009 0.01980 -0.0131 0.3597 1.0000 5.500 0.6807 0.03178 0.02181 -0.0128 0.3489 1.0000 5.750 0.7033 0.03333 0.02346 -0.0122 0.3394 1.0000 6.000 0.7262 0.03476 0.02495 -0.0116 0.3292 1.0000 6.250 0.7442 0.03705 0.02755 -0.0113 0.3201 1.0000 6.500 0.7700 0.03815 0.02850 -0.0103 0.3105 1.0000 6.750 0.7809 0.04127 0.03214 -0.0102 0.3011 1.0000 7.000 0.8020 0.04301 0.03390 -0.0094 0.2917 1.0000 7.250 0.8118 0.04627 0.03751 -0.0092 0.2829 1.0000 7.500 0.8279 0.04880 0.04014 -0.0086 0.2750 1.0000 7.750 0.8212 0.05433 0.04611 -0.0096 0.2697 1.0000 8.000 0.8548 0.05471 0.04631 -0.0077 0.2607 1.0000 8.250 0.8312 0.06233 0.05435 -0.0099 0.2592 1.0000 8.500 0.8070 0.07018 0.06236 -0.0128 0.2596 1.0000 8.750 0.7928 0.07724 0.06950 -0.0156 0.2613 1.0000 9.000 0.6054 0.10834 0.10022 -0.0486 0.4276 1.0000