XFOIL Version 6.94 Calculated polar for: NACA 63-210 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 0.1728 0.00673 0.00170 -0.0405 0.7313 0.7528 0.500 0.2280 0.00676 0.00168 -0.0402 0.7006 0.7739 1.000 0.2823 0.00684 0.00167 -0.0396 0.6559 0.7962 1.500 0.3358 0.00701 0.00175 -0.0389 0.5998 0.8198 2.000 0.3855 0.00761 0.00191 -0.0376 0.4660 0.8446 2.500 0.4230 0.01032 0.00289 -0.0355 0.0370 0.8707 3.000 0.4728 0.01064 0.00331 -0.0340 0.0313 0.8971 3.500 0.5198 0.01099 0.00383 -0.0318 0.0306 0.9268 4.000 0.5670 0.01142 0.00442 -0.0298 0.0305 0.9657 4.500 0.6292 0.01227 0.00528 -0.0318 0.0254 1.0000 5.000 0.6836 0.01314 0.00616 -0.0323 0.0218 1.0000 5.500 0.7373 0.01394 0.00699 -0.0325 0.0178 1.0000 6.000 0.7856 0.01548 0.00863 -0.0319 0.0155 1.0000 6.500 0.8475 0.01463 0.00762 -0.0333 0.0109 1.0000 7.000 0.8971 0.01580 0.00895 -0.0327 0.0035 1.0000 7.500 0.9460 0.01699 0.01016 -0.0320 0.0028 1.0000 8.000 0.9955 0.01797 0.01121 -0.0315 0.0025 1.0000 8.500 1.0422 0.01933 0.01277 -0.0305 0.0024 1.0000 9.000 1.0857 0.02109 0.01481 -0.0290 0.0023 1.0000 9.500 1.1256 0.02330 0.01741 -0.0273 0.0023 1.0000 10.000 1.1600 0.02629 0.02087 -0.0249 0.0023 1.0000 10.500 1.1848 0.03033 0.02549 -0.0218 0.0023 1.0000 11.000 1.1895 0.03582 0.03165 -0.0171 0.0024 1.0000 11.500 1.1616 0.04336 0.03988 -0.0113 0.0025 1.0000 12.000 1.1156 0.05382 0.05093 -0.0102 0.0025 1.0000 12.500 1.0612 0.06808 0.06566 -0.0158 0.0026 1.0000 14.000 0.8666 0.15270 0.15057 -0.0647 0.0031 1.0000 14.500 0.8590 0.16742 0.16525 -0.0723 0.0034 1.0000 15.000 0.8600 0.17993 0.17772 -0.0787 0.0036 1.0000 15.500 0.8672 0.19050 0.18828 -0.0843 0.0039 1.0000 16.000 0.7173 0.19099 0.18912 -0.0778 0.0041 1.0000