XFOIL Version 6.94 Calculated polar for: GOA 1b 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1767 0.02897 0.01524 -0.0311 1.0000 0.1671 -2.750 -0.1494 0.02708 0.01341 -0.0301 1.0000 0.1916 -2.500 -0.1228 0.02499 0.01167 -0.0290 1.0000 0.2387 -2.250 -0.0975 0.01983 0.00942 -0.0252 1.0000 1.0000 -2.000 -0.0709 0.01981 0.00836 -0.0238 1.0000 1.0000 -1.750 -0.0457 0.01982 0.00783 -0.0226 1.0000 1.0000 -1.500 -0.0210 0.01984 0.00742 -0.0216 1.0000 1.0000 -1.250 0.0035 0.01988 0.00709 -0.0207 1.0000 1.0000 -1.000 0.0279 0.01992 0.00684 -0.0198 1.0000 1.0000 -0.750 0.0522 0.01998 0.00665 -0.0189 1.0000 1.0000 -0.500 0.0766 0.02006 0.00661 -0.0180 1.0000 1.0000 -0.250 0.1011 0.02015 0.00671 -0.0171 1.0000 1.0000 0.000 0.1262 0.02025 0.00698 -0.0161 1.0000 1.0000 0.250 0.1529 0.02032 0.00750 -0.0150 1.0000 1.0000 0.750 0.2605 0.02476 0.00947 -0.0199 0.3392 1.0000 1.000 0.2877 0.02610 0.01051 -0.0191 0.3183 1.0000 1.250 0.3162 0.02748 0.01165 -0.0186 0.2989 1.0000 1.500 0.3457 0.02888 0.01293 -0.0183 0.2810 1.0000 1.750 0.3773 0.03034 0.01443 -0.0183 0.2713 1.0000 2.000 0.4107 0.03184 0.01618 -0.0186 0.2685 1.0000 2.250 0.4442 0.03347 0.01818 -0.0191 0.2678 1.0000 2.500 0.4775 0.03526 0.02039 -0.0196 0.2693 1.0000 2.750 0.5092 0.03729 0.02287 -0.0202 0.2724 1.0000 3.000 0.5403 0.03960 0.02556 -0.0208 0.2765 1.0000 3.250 0.5733 0.04200 0.02866 -0.0224 0.2860 1.0000 3.500 0.6046 0.04480 0.03201 -0.0240 0.2962 1.0000 3.750 0.6348 0.04819 0.03578 -0.0258 0.3092 1.0000 4.000 0.6680 0.05162 0.04003 -0.0305 0.3325 1.0000 4.250 0.6966 0.05564 0.04456 -0.0351 0.3526 1.0000 4.500 0.7226 0.06007 0.04936 -0.0399 0.3739 1.0000 5.000 0.7490 0.07248 0.06254 -0.0689 0.5056 1.0000