XFOIL Version 6.94 Calculated polar for: RAE 5214 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 0.2347 0.00712 0.00258 -0.0543 0.8244 0.6945 0.500 0.2897 0.00712 0.00248 -0.0534 0.7912 0.7034 1.000 0.3433 0.00719 0.00244 -0.0522 0.7428 0.7124 2.000 0.4345 0.00887 0.00265 -0.0474 0.3835 0.7307 2.500 0.4799 0.01027 0.00316 -0.0458 0.1888 0.7405 3.000 0.5312 0.01087 0.00356 -0.0449 0.1453 0.7501 3.500 0.5841 0.01135 0.00397 -0.0441 0.1255 0.7614 4.000 0.6361 0.01179 0.00442 -0.0432 0.1111 0.7731 4.500 0.6881 0.01227 0.00488 -0.0423 0.0983 0.7866 5.000 0.7390 0.01279 0.00545 -0.0413 0.0882 0.8040 5.500 0.7889 0.01329 0.00604 -0.0400 0.0793 0.8269 6.000 0.8355 0.01362 0.00658 -0.0379 0.0710 0.8755 6.500 0.8911 0.01408 0.00715 -0.0379 0.0624 1.0000 7.000 0.9417 0.01489 0.00790 -0.0371 0.0544 1.0000 8.000 1.0373 0.01700 0.01003 -0.0344 0.0433 1.0000 9.000 1.1266 0.01966 0.01281 -0.0309 0.0377 1.0000 9.500 1.1678 0.02152 0.01477 -0.0288 0.0360 1.0000 10.000 1.2090 0.02321 0.01667 -0.0267 0.0346 1.0000 10.500 1.2478 0.02504 0.01864 -0.0245 0.0334 1.0000 11.000 1.2831 0.02727 0.02097 -0.0220 0.0324 1.0000 11.500 1.3116 0.03005 0.02408 -0.0188 0.0315 1.0000 12.000 1.3317 0.03285 0.02725 -0.0145 0.0308 1.0000 12.500 1.3369 0.03598 0.03072 -0.0086 0.0303 1.0000 13.000 1.3329 0.03975 0.03486 -0.0030 0.0299 1.0000 13.500 1.3196 0.04459 0.04006 0.0012 0.0296 1.0000 14.000 1.2949 0.05118 0.04704 0.0033 0.0294 1.0000 14.500 1.2525 0.06122 0.05752 0.0013 0.0294 1.0000