XFOIL Version 6.94 Calculated polar for: GOE 142 (MVA H.19) 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 ------- -------- --------- --------- -------- ------- ------- 0.000 0.4035 0.00820 0.00180 -0.0768 0.7791 0.0901 0.500 0.4576 0.00799 0.00159 -0.0761 0.7330 0.1464 1.000 0.5070 0.00612 0.00159 -0.0748 0.6851 1.0000 1.500 0.5609 0.00646 0.00158 -0.0742 0.6329 1.0000 2.000 0.6147 0.00685 0.00166 -0.0735 0.5802 1.0000 2.500 0.6684 0.00727 0.00181 -0.0730 0.5283 1.0000 3.000 0.7221 0.00772 0.00202 -0.0725 0.4802 1.0000 3.500 0.7756 0.00819 0.00229 -0.0721 0.4355 1.0000 4.000 0.8290 0.00868 0.00262 -0.0717 0.3943 1.0000 4.500 0.8821 0.00920 0.00300 -0.0712 0.3568 1.0000 5.000 0.9350 0.00975 0.00345 -0.0707 0.3224 1.0000 5.500 0.9874 0.01034 0.00397 -0.0702 0.2910 1.0000 6.500 1.0908 0.01163 0.00513 -0.0691 0.2206 1.0000 7.000 1.1415 0.01236 0.00581 -0.0684 0.1872 1.0000 7.500 1.1908 0.01326 0.00662 -0.0676 0.1478 1.0000 8.000 1.2328 0.01520 0.00806 -0.0661 0.0598 1.0000 8.500 1.2720 0.01754 0.01028 -0.0638 0.0260 1.0000 9.000 1.3120 0.01951 0.01242 -0.0616 0.0193 1.0000 9.500 1.3446 0.02211 0.01528 -0.0585 0.0162 1.0000 10.000 1.3758 0.02453 0.01793 -0.0553 0.0143 1.0000 10.500 1.3932 0.02810 0.02172 -0.0506 0.0130 1.0000 11.000 1.4060 0.03157 0.02550 -0.0454 0.0120 1.0000 11.500 1.4147 0.03516 0.02942 -0.0403 0.0114 1.0000 12.000 1.4176 0.03963 0.03424 -0.0361 0.0109 1.0000 12.500 1.4152 0.04499 0.03994 -0.0331 0.0105 1.0000 13.000 1.4087 0.05112 0.04639 -0.0320 0.0100 1.0000 13.500 1.3977 0.05851 0.05404 -0.0329 0.0096 1.0000 14.000 1.3754 0.06833 0.06418 -0.0352 0.0092 1.0000 14.500 1.3448 0.08040 0.07661 -0.0399 0.0091 1.0000 15.000 1.3135 0.09371 0.09028 -0.0465 0.0091 1.0000 15.500 1.2823 0.10836 0.10524 -0.0548 0.0092 1.0000 16.000 1.2520 0.12427 0.12146 -0.0645 0.0093 1.0000 16.500 1.2114 0.14480 0.14236 -0.0776 0.0096 1.0000