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 = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1750 0.02766 0.01460 -0.0311 1.0000 0.1380 -2.750 -0.1467 0.02603 0.01271 -0.0298 1.0000 0.1479 -2.500 -0.1206 0.02427 0.01115 -0.0286 1.0000 0.1718 -2.250 -0.0947 0.02175 0.00951 -0.0279 1.0000 0.2664 -2.000 -0.0704 0.01847 0.00804 -0.0242 1.0000 1.0000 -1.750 -0.0449 0.01848 0.00748 -0.0230 1.0000 1.0000 -1.500 -0.0200 0.01851 0.00712 -0.0219 1.0000 1.0000 -1.250 0.0046 0.01855 0.00682 -0.0209 1.0000 1.0000 -1.000 0.0290 0.01860 0.00658 -0.0200 1.0000 1.0000 -0.750 0.0535 0.01867 0.00641 -0.0191 1.0000 1.0000 -0.500 0.0780 0.01875 0.00640 -0.0182 1.0000 1.0000 -0.250 0.1028 0.01883 0.00654 -0.0173 1.0000 1.0000 0.000 0.1286 0.01892 0.00689 -0.0162 1.0000 1.0000 0.250 0.1577 0.01894 0.00754 -0.0151 1.0000 1.0000 0.500 0.2339 0.02292 0.00842 -0.0211 0.3131 1.0000 0.750 0.2592 0.02420 0.00929 -0.0201 0.2909 1.0000 1.000 0.2864 0.02540 0.01025 -0.0194 0.2699 1.0000 1.250 0.3150 0.02665 0.01132 -0.0189 0.2533 1.0000 1.500 0.3461 0.02795 0.01261 -0.0188 0.2462 1.0000 1.750 0.3786 0.02945 0.01414 -0.0189 0.2430 1.0000 2.000 0.4119 0.03089 0.01587 -0.0191 0.2421 1.0000 2.250 0.4442 0.03252 0.01782 -0.0193 0.2427 1.0000 2.500 0.4761 0.03413 0.01993 -0.0195 0.2455 1.0000 2.750 0.5084 0.03609 0.02245 -0.0201 0.2510 1.0000 3.000 0.5389 0.03847 0.02519 -0.0206 0.2570 1.0000 3.250 0.5712 0.04089 0.02826 -0.0219 0.2677 1.0000 3.500 0.6030 0.04374 0.03169 -0.0235 0.2814 1.0000 3.750 0.6340 0.04693 0.03542 -0.0257 0.2970 1.0000 4.000 0.6617 0.05052 0.03937 -0.0277 0.3077 1.0000 4.250 0.6930 0.05439 0.04393 -0.0329 0.3327 1.0000 4.500 0.7283 0.05987 0.05006 -0.0428 0.3917 1.0000 5.000 0.6917 0.07448 0.06505 -0.0861 0.6510 1.0000