XFOIL Version 6.94 Calculated polar for: NACA 66,2-(1.8)15.5 A=.6 P-51 RO 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 0.1298 0.01129 0.00653 -0.0267 0.7612 0.8110 0.500 0.1897 0.01094 0.00609 -0.0275 0.7547 0.8141 1.000 0.2448 0.01087 0.00604 -0.0279 0.7492 0.8172 1.500 0.2994 0.01072 0.00592 -0.0282 0.7428 0.8206 2.000 0.3569 0.01051 0.00571 -0.0289 0.7369 0.8245 2.500 0.4156 0.01026 0.00543 -0.0297 0.7288 0.8277 3.000 0.4707 0.00972 0.00492 -0.0297 0.7127 0.8309 3.500 0.5234 0.00941 0.00470 -0.0291 0.6940 0.8346 4.000 0.5717 0.00915 0.00446 -0.0276 0.6515 0.8389 4.500 0.5653 0.01016 0.00460 -0.0158 0.4673 0.8463 5.000 0.5292 0.01171 0.00558 0.0007 0.3347 0.8558 5.500 0.5071 0.01397 0.00711 0.0128 0.1697 0.8665 6.000 0.5059 0.01600 0.00850 0.0209 0.0416 0.8768 6.500 0.5357 0.01692 0.00946 0.0244 0.0361 0.8868 7.000 0.5667 0.01778 0.01042 0.0275 0.0340 0.8971 7.500 0.5946 0.01882 0.01152 0.0310 0.0323 0.9101 8.000 0.6185 0.02001 0.01283 0.0351 0.0312 0.9263 8.500 0.6495 0.02115 0.01410 0.0377 0.0298 0.9470 9.000 0.6970 0.02284 0.01585 0.0362 0.0284 0.9672 9.500 0.7546 0.02478 0.01784 0.0327 0.0271 0.9826 10.000 0.8074 0.02640 0.01954 0.0309 0.0259 1.0000 10.500 0.8418 0.02771 0.02084 0.0328 0.0245 1.0000 11.000 0.8893 0.02939 0.02259 0.0334 0.0234 1.0000 11.500 0.9347 0.03142 0.02485 0.0341 0.0227 1.0000 12.000 0.9790 0.03409 0.02780 0.0348 0.0220 1.0000 12.500 1.0123 0.03767 0.03175 0.0364 0.0217 1.0000 13.000 1.0246 0.04373 0.03844 0.0398 0.0224 1.0000 13.500 1.0153 0.05001 0.04519 0.0442 0.0232 1.0000 14.000 1.0124 0.05655 0.05200 0.0468 0.0237 1.0000