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 = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -4.000 -0.4529 0.01189 0.00490 0.0192 0.6031 0.0240 -3.500 -0.4013 0.01144 0.00442 0.0201 0.5953 0.0235 -3.000 -0.3527 0.01069 0.00362 0.0216 0.5880 0.0231 -2.500 -0.3011 0.01017 0.00306 0.0226 0.5786 0.0229 -2.000 -0.2484 0.00976 0.00259 0.0234 0.5601 0.0226 -1.500 -0.1996 0.01012 0.00239 0.0245 0.3680 0.0222 -1.000 -0.1481 0.01029 0.00224 0.0253 0.2747 0.0218 -0.500 -0.0945 0.01025 0.00213 0.0257 0.2562 0.0215 0.000 -0.0405 0.01020 0.00204 0.0261 0.2404 0.0213 0.500 0.0112 0.01049 0.00201 0.0268 0.1480 0.0214 1.000 0.0626 0.01084 0.00211 0.0275 0.0839 0.0221 1.500 0.1154 0.01082 0.00222 0.0281 0.0801 0.0677 2.000 0.1569 0.00981 0.00218 0.0303 0.0769 0.4098 2.500 0.1958 0.00904 0.00236 0.0334 0.0721 0.6674 3.000 0.2406 0.00877 0.00253 0.0358 0.0708 0.7862 3.500 0.2883 0.00874 0.00279 0.0376 0.0682 0.8624 4.000 0.3401 0.00892 0.00307 0.0386 0.0649 0.9035 4.500 0.3948 0.00926 0.00346 0.0387 0.0612 0.9315 5.000 0.4537 0.00945 0.00366 0.0380 0.0593 0.9506 5.500 0.5177 0.00975 0.00393 0.0360 0.0554 0.9629 6.000 0.5806 0.01015 0.00428 0.0342 0.0514 0.9726 6.500 0.6514 0.01057 0.00468 0.0306 0.0473 0.9763 7.000 0.7121 0.01132 0.00523 0.0289 0.0241 0.9832 7.500 0.7805 0.01205 0.00598 0.0255 0.0201 0.9863 8.000 0.8454 0.01276 0.00675 0.0229 0.0191 0.9908 8.500 0.9107 0.01350 0.00757 0.0201 0.0184 0.9949 9.000 0.9769 0.01440 0.00855 0.0168 0.0176 0.9984 9.500 1.0272 0.01534 0.00959 0.0168 0.0172 1.0000 10.000 1.0563 0.01632 0.01067 0.0210 0.0166 1.0000