XFOIL Version 6.94 Calculated polar for: NACA 66(2)-215 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 -0.1236 0.02088 0.01594 0.0111 0.7798 0.8077 0.500 -0.0825 0.02227 0.01743 0.0158 0.7725 0.8265 1.000 -0.0221 0.02250 0.01765 0.0153 0.7674 0.8364 1.500 0.0370 0.02226 0.01741 0.0140 0.7634 0.8406 2.000 0.1039 0.02192 0.01707 0.0114 0.7600 0.8435 2.500 0.2304 0.01986 0.01499 -0.0004 0.7577 0.8433 3.000 0.4009 0.01655 0.01165 -0.0195 0.7463 0.8413 3.500 0.4440 0.01550 0.01068 -0.0170 0.7314 0.8446 4.000 0.5500 0.01344 0.00855 -0.0254 0.7059 0.8453 4.500 0.5794 0.01246 0.00730 -0.0194 0.6010 0.8481 5.000 0.5310 0.01249 0.00656 0.0003 0.4585 0.8583 5.500 0.4833 0.01399 0.00757 0.0183 0.3382 0.8692 6.000 0.4469 0.01616 0.00902 0.0324 0.1752 0.8805 6.500 0.4386 0.01807 0.01029 0.0413 0.0570 0.8904 7.000 0.4629 0.01896 0.01131 0.0454 0.0467 0.8970 7.500 0.4850 0.01990 0.01260 0.0493 0.0408 0.9076 8.000 0.4993 0.02143 0.01431 0.0547 0.0324 0.9216 8.500 0.5199 0.02335 0.01641 0.0592 0.0284 0.9344 9.000 0.5717 0.02532 0.01858 0.0598 0.0269 0.9435 10.000 0.7102 0.03076 0.02458 0.0545 0.0285 0.9608 10.500 0.7696 0.03657 0.03114 0.0522 0.0315 0.9696 11.000 0.7988 0.04025 0.03496 0.0511 0.0274 0.9826 11.500 0.7997 0.04760 0.04280 0.0523 0.0256 0.9957 12.000 0.7670 0.05143 0.04688 0.0597 0.0253 1.0000 12.500 0.7343 0.05654 0.05224 0.0651 0.0252 1.0000 13.000 0.7095 0.06140 0.05732 0.0685 0.0253 1.0000 13.500 0.6838 0.06688 0.06301 0.0704 0.0256 1.0000 14.000 0.6514 0.07396 0.07030 0.0704 0.0260 1.0000 14.500 0.6151 0.08306 0.07962 0.0676 0.0267 1.0000