XFOIL Version 6.94 Calculated polar for: GOE 506 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.500 0.4179 0.01056 0.00330 -0.0689 0.5113 0.3750 1.000 0.4283 0.01107 0.00380 -0.0602 0.4604 0.4520 1.500 0.4364 0.01096 0.00428 -0.0510 0.4210 0.6660 3.000 0.7358 0.01285 0.00631 -0.0821 0.3497 1.0000 3.500 0.7635 0.01341 0.00674 -0.0772 0.3406 1.0000 4.000 0.7993 0.01377 0.00712 -0.0738 0.3353 1.0000 4.500 0.8272 0.01446 0.00762 -0.0693 0.3158 1.0000 5.000 0.8584 0.01509 0.00809 -0.0655 0.2897 1.0000 5.500 0.8920 0.01577 0.00862 -0.0622 0.2651 1.0000 6.000 0.9045 0.01758 0.00980 -0.0561 0.1697 1.0000 6.500 0.9319 0.01885 0.01091 -0.0524 0.1464 1.0000 7.000 0.9638 0.01992 0.01192 -0.0494 0.1319 1.0000 7.500 0.9684 0.02268 0.01418 -0.0432 0.0280 1.0000 8.000 0.9971 0.02411 0.01559 -0.0402 0.0039 1.0000 8.500 1.0284 0.02539 0.01693 -0.0377 0.0038 1.0000 9.000 1.0591 0.02678 0.01841 -0.0352 0.0040 1.0000 9.500 1.0869 0.02841 0.02015 -0.0327 0.0041 1.0000 10.000 1.1122 0.03029 0.02216 -0.0301 0.0043 1.0000 10.500 1.1351 0.03244 0.02445 -0.0274 0.0046 1.0000 11.000 1.1562 0.03483 0.02698 -0.0249 0.0050 1.0000 11.500 1.1716 0.03778 0.03011 -0.0223 0.0054 1.0000 12.000 1.1788 0.04165 0.03419 -0.0196 0.0058 1.0000 12.500 1.1868 0.04570 0.03842 -0.0174 0.0064 1.0000 13.000 1.1867 0.05086 0.04380 -0.0155 0.0071 1.0000 13.500 1.1742 0.05785 0.05104 -0.0143 0.0075 1.0000 14.000 1.1511 0.06689 0.06035 -0.0145 0.0077 1.0000 14.500 1.1478 0.07399 0.06766 -0.0152 0.0084 1.0000 15.000 1.1242 0.08402 0.07795 -0.0167 0.0091 1.0000 15.500 1.0997 0.09435 0.08849 -0.0187 0.0095 1.0000 16.000 1.0856 0.10320 0.09748 -0.0205 0.0102 1.0000 16.500 1.0859 0.10868 0.10296 -0.0199 0.0122 1.0000