XFOIL Version 6.96 Calculated polar for: RAF-48 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -17.250 -0.7767 0.10552 0.10252 -0.0428 1.0000 0.0248 -17.000 -0.8387 0.08965 0.08639 -0.0532 1.0000 0.0248 -16.750 -0.8747 0.07966 0.07622 -0.0597 1.0000 0.0247 -16.500 -0.9027 0.07124 0.06762 -0.0654 1.0000 0.0247 -16.250 -0.9261 0.06385 0.06005 -0.0707 1.0000 0.0248 -16.000 -0.9496 0.05703 0.05304 -0.0750 1.0000 0.0250 -15.750 -0.9616 0.05219 0.04808 -0.0779 1.0000 0.0252 -15.500 -0.9686 0.04826 0.04406 -0.0799 1.0000 0.0254 -15.250 -0.9710 0.04513 0.04087 -0.0815 1.0000 0.0257 -15.000 -0.9745 0.04208 0.03773 -0.0826 1.0000 0.0259 -14.750 -0.9753 0.03959 0.03515 -0.0832 1.0000 0.0262 -14.500 -0.9747 0.03749 0.03297 -0.0833 1.0000 0.0267 -14.250 -0.9752 0.03545 0.03084 -0.0829 1.0000 0.0271 -14.000 -0.9763 0.03354 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0.04436 0.0073 0.0398 1.0000 15.250 1.2957 0.05336 0.04734 0.0073 0.0387 1.0000 15.500 1.2902 0.05681 0.05084 0.0072 0.0375 1.0000 15.750 1.2829 0.06054 0.05463 0.0069 0.0365 1.0000 16.000 1.2843 0.06333 0.05752 0.0066 0.0357 1.0000 16.250 1.2848 0.06629 0.06056 0.0062 0.0344 1.0000 16.500 1.2822 0.06967 0.06403 0.0057 0.0337 1.0000 16.750 1.2786 0.07321 0.06763 0.0051 0.0329 1.0000 17.000 1.2717 0.07722 0.07168 0.0043 0.0320 1.0000 17.250 1.2639 0.08135 0.07587 0.0035 0.0312 1.0000 17.500 1.2642 0.08451 0.07912 0.0028 0.0307 1.0000 17.750 1.2625 0.08801 0.08272 0.0019 0.0298 1.0000 18.000 1.2603 0.09156 0.08635 0.0010 0.0292 1.0000 18.250 1.2575 0.09524 0.09008 0.0000 0.0284 1.0000 18.500 1.2540 0.09906 0.09395 -0.0012 0.0277 1.0000 18.750 1.2496 0.10286 0.09778 -0.0022 0.0272 1.0000