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 = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -4.500 -0.3801 0.02153 0.01192 -0.0144 0.8796 0.0739 -4.000 -0.3430 0.02010 0.01048 -0.0118 0.8743 0.0806 -3.500 -0.3063 0.01865 0.00922 -0.0089 0.8695 0.1199 -3.000 -0.3000 0.01617 0.00924 -0.0003 0.8639 0.6272 -2.500 -0.2544 0.01650 0.00971 0.0030 0.8582 0.7159 -2.000 -0.2033 0.01703 0.01027 0.0052 0.8522 0.7715 -1.500 -0.1424 0.01800 0.01126 0.0068 0.8446 0.8254 -1.000 -0.0686 0.01892 0.01207 0.0066 0.8295 0.8599 -0.500 0.0564 0.02045 0.01343 -0.0003 0.8058 0.8990 0.000 0.1243 0.02036 0.01325 -0.0019 0.7861 0.9180 0.500 0.1878 0.01994 0.01267 -0.0020 0.7546 0.9348 1.000 0.2524 0.01944 0.01212 -0.0031 0.7268 0.9522 1.500 0.3252 0.01866 0.01128 -0.0056 0.6901 0.9704 2.000 0.4011 0.01743 0.01005 -0.0091 0.6364 0.9878 2.500 0.4634 0.01609 0.00821 -0.0105 0.4364 1.0000 3.000 0.4924 0.01697 0.00851 -0.0072 0.3609 1.0000 3.500 0.5240 0.01758 0.00889 -0.0044 0.2920 1.0000 4.000 0.5535 0.01853 0.00942 -0.0008 0.2020 1.0000 4.500 0.5848 0.01936 0.00988 0.0027 0.1733 1.0000 5.000 0.6172 0.02039 0.01074 0.0063 0.1592 1.0000 5.500 0.6508 0.02157 0.01182 0.0096 0.1481 1.0000 6.000 0.6859 0.02273 0.01292 0.0127 0.1391 1.0000 6.500 0.7239 0.02422 0.01445 0.0154 0.1348 1.0000 7.000 0.7629 0.02569 0.01613 0.0181 0.1314 1.0000 7.500 0.8011 0.02716 0.01775 0.0207 0.1268 1.0000 8.000 0.8409 0.02901 0.01975 0.0229 0.1243 1.0000 8.500 0.8812 0.03149 0.02235 0.0248 0.1221 1.0000 9.000 0.9142 0.03371 0.02508 0.0277 0.1200 1.0000 9.500 0.9438 0.03556 0.02736 0.0308 0.1147 1.0000 10.000 0.9783 0.03731 0.02925 0.0332 0.1099 1.0000