XFOIL Version 6.94 Calculated polar for: manu4/26 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0032 0.10525 0.08788 -0.0162 0.9999 1.0001 -2.750 0.0058 0.10321 0.08619 -0.0155 0.9999 1.0001 -2.500 0.0078 0.10117 0.08450 -0.0148 0.9999 1.0001 -2.250 0.0090 0.09914 0.08284 -0.0141 0.9999 1.0001 -2.000 0.0094 0.09714 0.08123 -0.0132 0.9999 1.0001 -1.750 0.0087 0.09515 0.07967 -0.0122 0.9999 1.0001 -1.500 0.0065 0.09318 0.07813 -0.0111 0.9999 1.0001 -1.250 0.0026 0.09129 0.07670 -0.0098 0.9999 1.0001 -1.000 -0.0036 0.08945 0.07535 -0.0083 0.9999 1.0001 -0.750 -0.0119 0.08781 0.07419 -0.0066 0.9999 1.0001 -0.500 -0.0217 0.08656 0.07340 -0.0050 0.9999 1.0001 -0.250 -0.0329 0.08603 0.07325 -0.0039 0.9999 1.0001 0.000 -0.0473 0.08679 0.07429 -0.0030 0.9999 1.0001 0.250 -0.0652 0.08865 0.07623 -0.0023 0.9999 1.0001 0.500 -0.0791 0.09072 0.07816 -0.0023 0.9999 1.0001 0.750 -0.0855 0.09268 0.07985 -0.0034 0.9999 1.0001 1.000 -0.0847 0.09465 0.08146 -0.0055 0.9999 1.0001 1.250 -0.0773 0.09671 0.08307 -0.0085 0.9999 1.0001 1.500 -0.0650 0.09897 0.08480 -0.0123 0.9999 1.0001 1.750 -0.0480 0.10143 0.08668 -0.0167 0.9999 1.0001 2.000 -0.0277 0.10412 0.08874 -0.0215 0.9999 1.0001 2.250 -0.0047 0.10700 0.09094 -0.0267 0.9999 1.0001 2.500 0.0206 0.11008 0.09330 -0.0321 0.9999 1.0001 2.750 0.0472 0.11333 0.09577 -0.0375 0.9999 1.0001 3.000 0.0743 0.11668 0.09831 -0.0427 0.9999 1.0001 3.250 0.1008 0.12006 0.10084 -0.0475 0.9999 1.0001 3.500 0.1259 0.12342 0.10338 -0.0518 0.9999 1.0001 3.750 0.1495 0.12673 0.10588 -0.0555 0.9999 1.0001 4.000 0.1713 0.12992 0.10831 -0.0586 0.9999 1.0001 4.250 0.1915 0.13306 0.11071 -0.0613 0.9999 1.0001 4.500 0.2106 0.13607 0.11302 -0.0636 0.9999 1.0001 4.750 0.2284 0.13903 0.11529 -0.0656 0.9999 1.0001 5.000 0.2453 0.14192 0.11755 -0.0673 0.9999 1.0001