XFOIL Version 6.94 Calculated polar for: WBL FX 71-089A AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0576 0.04679 0.02551 -0.0129 0.9989 1.0011 -2.750 -0.0321 0.04671 0.02506 -0.0135 0.9989 1.0011 -2.500 -0.0068 0.04674 0.02482 -0.0138 0.9989 1.0011 -2.250 0.0182 0.04685 0.02475 -0.0140 0.9989 1.0011 -2.000 0.0429 0.04704 0.02483 -0.0140 0.9989 1.0011 -1.750 0.0671 0.04731 0.02506 -0.0139 0.9989 1.0011 -1.500 0.0910 0.04764 0.02542 -0.0138 0.9989 1.0011 -1.250 0.1144 0.04806 0.02592 -0.0137 0.9989 1.0011 -1.000 0.1373 0.04857 0.02657 -0.0135 0.9989 1.0011 -0.750 0.1596 0.04917 0.02739 -0.0133 0.9989 1.0011 -0.500 0.1814 0.04991 0.02841 -0.0132 0.9989 1.0011 -0.250 0.2022 0.05083 0.02969 -0.0132 0.9989 1.0011 0.000 0.2212 0.05204 0.03133 -0.0133 0.9989 1.0011 0.250 0.2363 0.05383 0.03364 -0.0138 0.9989 1.0011 0.500 0.2424 0.05696 0.03722 -0.0152 0.9989 1.0011 0.750 0.2350 0.06202 0.04244 -0.0179 0.9989 1.0011 1.000 0.2273 0.06728 0.04764 -0.0212 0.9989 1.0011 1.250 0.2256 0.07180 0.05208 -0.0241 0.9989 1.0011 1.500 0.2279 0.07582 0.05602 -0.0267 0.9989 1.0011 1.750 0.2327 0.07955 0.05968 -0.0290 0.9989 1.0011 2.000 0.2394 0.08316 0.06321 -0.0313 0.9989 1.0011 2.250 0.2471 0.08665 0.06663 -0.0333 0.9989 1.0011 2.500 0.2556 0.09007 0.06998 -0.0353 0.9989 1.0011 2.750 0.2647 0.09345 0.07329 -0.0372 0.9989 1.0011 3.000 0.2742 0.09680 0.07658 -0.0391 0.9989 1.0011 3.250 0.2841 0.10015 0.07985 -0.0408 0.9989 1.0011 3.500 0.2943 0.10348 0.08312 -0.0426 0.9989 1.0011 3.750 0.3047 0.10680 0.08638 -0.0443 0.9989 1.0011 4.000 0.3154 0.11013 0.08964 -0.0461 0.9989 1.0011 4.250 0.3262 0.11346 0.09291 -0.0478 0.9989 1.0011 4.500 0.3371 0.11679 0.09618 -0.0495 0.9989 1.0011 4.750 0.3483 0.12012 0.09945 -0.0511 0.9989 1.0011 5.000 0.3595 0.12346 0.10273 -0.0528 0.9989 1.0011