XFOIL Version 6.94 Calculated polar for: NACA 63-412 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3320 0.00928 0.00368 -0.0769 0.7498 0.7455 0.500 0.3872 0.00932 0.00374 -0.0766 0.7342 0.7629 1.000 0.4424 0.00940 0.00382 -0.0763 0.7192 0.7810 1.500 0.4975 0.00949 0.00389 -0.0759 0.7040 0.7996 2.000 0.5517 0.00957 0.00406 -0.0754 0.6875 0.8184 2.500 0.6048 0.00969 0.00425 -0.0746 0.6713 0.8376 3.000 0.6570 0.00978 0.00440 -0.0735 0.6535 0.8568 3.500 0.7075 0.00981 0.00450 -0.0721 0.6314 0.8776 4.000 0.7551 0.00978 0.00457 -0.0699 0.6045 0.8998 4.500 0.7972 0.00971 0.00451 -0.0666 0.5669 0.9279 5.000 0.8484 0.00975 0.00461 -0.0654 0.5288 0.9725 5.500 0.9063 0.01023 0.00496 -0.0666 0.4646 1.0000 6.000 0.9427 0.01196 0.00583 -0.0643 0.2885 1.0000 6.500 0.9654 0.01501 0.00769 -0.0605 0.0960 1.0000 7.000 0.9972 0.01711 0.00948 -0.0573 0.0510 1.0000 7.500 1.0293 0.01888 0.01120 -0.0541 0.0386 1.0000 8.500 1.0815 0.02217 0.01457 -0.0460 0.0273 1.0000 9.000 1.1029 0.02411 0.01655 -0.0419 0.0245 1.0000 9.500 1.1269 0.02614 0.01872 -0.0384 0.0226 1.0000 10.000 1.1490 0.02873 0.02131 -0.0352 0.0213 1.0000 10.500 1.1760 0.03123 0.02405 -0.0327 0.0203 1.0000 11.000 1.2024 0.03389 0.02691 -0.0304 0.0192 1.0000 11.500 1.2248 0.03666 0.02973 -0.0284 0.0179 1.0000 12.000 1.2395 0.03959 0.03304 -0.0263 0.0164 1.0000 12.500 1.2536 0.04267 0.03623 -0.0249 0.0153 1.0000 13.000 1.2668 0.04688 0.04067 -0.0236 0.0147 1.0000 13.500 1.2729 0.05188 0.04609 -0.0227 0.0143 1.0000 14.000 1.2732 0.05785 0.05245 -0.0226 0.0139 1.0000 14.500 1.2679 0.06477 0.05976 -0.0235 0.0135 1.0000 15.000 1.2581 0.07279 0.06812 -0.0258 0.0131 1.0000 15.500 1.2423 0.08251 0.07821 -0.0295 0.0129 1.0000 16.000 1.2203 0.09426 0.09034 -0.0352 0.0128 1.0000 16.500 1.1800 0.11115 0.10771 -0.0448 0.0129 1.0000 17.000 1.0252 0.16435 0.16173 -0.0785 0.0152 1.0000