XFOIL Version 6.96 Calculated polar for: BOEING 707 .19 SPAN AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -4.500 -0.4884 0.01464 0.00677 0.0149 0.6566 0.0289 -4.000 -0.4421 0.01375 0.00583 0.0168 0.6497 0.0286 -3.500 -0.3948 0.01300 0.00500 0.0185 0.6445 0.0283 -3.000 -0.3451 0.01239 0.00434 0.0198 0.6366 0.0280 -2.500 -0.2942 0.01193 0.00377 0.0210 0.6261 0.0278 -2.000 -0.2415 0.01151 0.00334 0.0218 0.6177 0.0277 -1.500 -0.1884 0.01121 0.00297 0.0226 0.6076 0.0278 -1.000 -0.1344 0.01088 0.00266 0.0232 0.5953 0.0283 0.000 -0.0326 0.01083 0.00220 0.0251 0.3800 0.0749 1.000 0.0480 0.00934 0.00202 0.0299 0.2670 0.5719 1.500 0.0939 0.00914 0.00217 0.0320 0.2470 0.6803 2.000 0.1401 0.00897 0.00233 0.0342 0.2150 0.7756 2.500 0.1811 0.00959 0.00281 0.0373 0.0965 0.8741 3.000 0.2341 0.01005 0.00330 0.0381 0.0863 0.9147 3.500 0.2936 0.01053 0.00382 0.0374 0.0817 0.9379 4.000 0.3576 0.01112 0.00442 0.0355 0.0778 0.9524 4.500 0.4230 0.01195 0.00523 0.0330 0.0736 0.9631 5.000 0.4940 0.01263 0.00591 0.0293 0.0709 0.9698 5.500 0.5666 0.01307 0.00636 0.0254 0.0671 0.9761 6.000 0.6357 0.01363 0.00691 0.0220 0.0638 0.9838 6.500 0.7069 0.01441 0.00770 0.0179 0.0611 0.9891 7.000 0.7789 0.01475 0.00810 0.0140 0.0595 0.9948 7.500 0.8549 0.01505 0.00846 0.0091 0.0566 0.9998 8.000 0.8953 0.01528 0.00872 0.0115 0.0539 1.0000 8.500 0.9326 0.01567 0.00912 0.0145 0.0513 1.0000 9.000 0.9694 0.01613 0.00967 0.0176 0.0486 1.0000 9.500 1.0068 0.01657 0.01014 0.0206 0.0425 1.0000 10.000 1.0334 0.01786 0.01125 0.0248 0.0238 1.0000