XFOIL Version 6.94 Calculated polar for: GOA 3 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1525 0.03085 0.01619 -0.0265 0.9999 1.0001 -2.750 -0.1268 0.03077 0.01510 -0.0258 0.9999 1.0001 -2.500 -0.1024 0.03071 0.01427 -0.0250 0.9999 1.0001 -2.250 -0.0783 0.03069 0.01359 -0.0241 0.9999 1.0001 -2.000 -0.0545 0.03068 0.01304 -0.0232 0.9999 1.0001 -1.750 -0.0307 0.03070 0.01255 -0.0223 0.9999 1.0001 -1.500 -0.0070 0.03074 0.01223 -0.0213 0.9999 1.0001 -1.250 0.0167 0.03081 0.01201 -0.0204 0.9999 1.0001 -1.000 0.0404 0.03089 0.01191 -0.0195 0.9999 1.0001 -0.750 0.0643 0.03099 0.01192 -0.0185 0.9999 1.0001 -0.500 0.0885 0.03111 0.01206 -0.0176 0.9999 1.0001 -0.250 0.1132 0.03124 0.01229 -0.0166 0.9999 1.0001 0.000 0.1389 0.03136 0.01275 -0.0156 0.9999 1.0001 0.250 0.1669 0.03145 0.01340 -0.0147 0.9999 1.0001 0.500 0.2001 0.03144 0.01429 -0.0145 0.9999 1.0001 0.750 0.2399 0.03143 0.01549 -0.0161 0.9999 1.0001 1.000 0.3835 0.03441 0.01710 -0.0329 0.5718 1.0001 1.250 0.4156 0.03675 0.01878 -0.0330 0.5363 1.0001 1.500 0.4483 0.03897 0.02073 -0.0337 0.5148 1.0001 1.750 0.4817 0.04113 0.02291 -0.0350 0.5008 1.0001 2.000 0.5150 0.04341 0.02531 -0.0365 0.4920 1.0001 2.250 0.5479 0.04578 0.02798 -0.0387 0.4870 1.0001 2.500 0.5798 0.04837 0.03091 -0.0412 0.4849 1.0001 2.750 0.6095 0.05119 0.03407 -0.0438 0.4847 1.0001 3.000 0.6361 0.05425 0.03744 -0.0465 0.4842 1.0001 3.250 0.6594 0.05753 0.04103 -0.0490 0.4837 1.0001 3.500 0.6792 0.06103 0.04478 -0.0516 0.4838 1.0001 3.750 0.6948 0.06476 0.04873 -0.0544 0.4858 1.0001 4.000 0.7079 0.06872 0.05284 -0.0572 0.4897 1.0001 4.250 0.7220 0.07282 0.05707 -0.0598 0.4946 1.0001 4.500 0.7203 0.07717 0.06153 -0.0637 0.5062 1.0001 4.750 0.7204 0.08154 0.06592 -0.0671 0.5184 1.0001 5.000 0.7195 0.08596 0.07031 -0.0704 0.5329 1.0001