XFOIL Version 6.96 Calculated polar for: BOEING 737 MIDSPAN 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 ------ -------- --------- --------- -------- -------- -------- 0.250 0.1572 0.00722 0.00152 -0.0176 0.6543 0.6792 0.500 0.1787 0.00679 0.00163 -0.0161 0.6502 0.8066 0.750 0.2013 0.00673 0.00177 -0.0145 0.6445 0.8827 1.000 0.2259 0.00679 0.00184 -0.0135 0.6380 0.9119 1.250 0.2501 0.00683 0.00194 -0.0122 0.6317 0.9415 1.500 0.2812 0.00698 0.00205 -0.0124 0.6264 0.9648 1.750 0.3152 0.00703 0.00212 -0.0137 0.6209 0.9722 2.000 0.3516 0.00711 0.00216 -0.0156 0.6145 0.9778 2.250 0.3869 0.00719 0.00223 -0.0173 0.6085 0.9832 2.500 0.4312 0.00726 0.00229 -0.0210 0.6006 0.9875 2.750 0.4721 0.00734 0.00237 -0.0239 0.5909 0.9932 3.000 0.5091 0.00741 0.00241 -0.0260 0.5769 0.9962 3.250 0.5455 0.00747 0.00246 -0.0281 0.5659 0.9991 3.500 0.5742 0.00751 0.00252 -0.0285 0.5537 1.0000 3.750 0.5991 0.00748 0.00254 -0.0280 0.5375 1.0000 4.000 0.6239 0.00740 0.00253 -0.0275 0.4954 1.0000 4.250 0.6450 0.00773 0.00251 -0.0264 0.4433 1.0000 4.500 0.6679 0.00803 0.00274 -0.0257 0.4085 1.0000 4.750 0.6894 0.00843 0.00293 -0.0247 0.3555 1.0000 5.000 0.7104 0.00886 0.00319 -0.0237 0.3104 1.0000 5.250 0.7312 0.00931 0.00345 -0.0227 0.2657 1.0000 5.500 0.7514 0.00982 0.00376 -0.0216 0.2210 1.0000 5.750 0.7724 0.01026 0.00407 -0.0206 0.1882 1.0000 6.000 0.7938 0.01068 0.00438 -0.0196 0.1622 1.0000 6.250 0.8154 0.01111 0.00471 -0.0187 0.1370 1.0000 6.500 0.8374 0.01155 0.00507 -0.0179 0.1126 1.0000 6.750 0.8594 0.01204 0.00549 -0.0171 0.0916 1.0000 7.000 0.8809 0.01262 0.00596 -0.0163 0.0704 1.0000 7.250 0.9020 0.01329 0.00652 -0.0154 0.0482 1.0000 7.500 0.9222 0.01411 0.00730 -0.0143 0.0353 1.0000 7.750 0.9445 0.01467 0.00793 -0.0135 0.0306 1.0000 8.000 0.9636 0.01563 0.00893 -0.0123 0.0260 1.0000 8.250 0.9866 0.01609 0.00947 -0.0117 0.0239 1.0000 8.500 1.0080 0.01673 0.01017 -0.0110 0.0217 1.0000 8.750 1.0242 0.01799 0.01150 -0.0094 0.0191 1.0000 9.000 1.0442 0.01877 0.01241 -0.0085 0.0176 1.0000 9.250 1.0658 0.01935 0.01305 -0.0078 0.0156 1.0000 9.500 1.0839 0.02027 0.01402 -0.0068 0.0139 1.0000 9.750 1.0934 0.02227 0.01618 -0.0045 0.0126 1.0000 10.000 1.1115 0.02320 0.01724 -0.0033 0.0118 1.0000 10.250 1.1272 0.02436 0.01852 -0.0020 0.0110 1.0000 10.500 1.1418 0.02555 0.01984 -0.0006 0.0103 1.0000 10.750 1.1547 0.02682 0.02121 0.0009 0.0098 1.0000 11.000 1.1626 0.02841 0.02292 0.0031 0.0094 1.0000 11.250 1.1631 0.03101 0.02571 0.0058 0.0090 1.0000 11.750 1.1554 0.03719 0.03242 0.0110 0.0086 1.0000 12.000 1.1554 0.03935 0.03480 0.0125 0.0085 1.0000 12.250 1.1531 0.04189 0.03755 0.0136 0.0084 1.0000 12.500 1.1489 0.04483 0.04069 0.0141 0.0083 1.0000 12.750 1.1421 0.04835 0.04442 0.0138 0.0081 1.0000 13.000 1.1331 0.05243 0.04872 0.0128 0.0080 1.0000 13.250 1.1216 0.05716 0.05365 0.0112 0.0080 1.0000 13.500 1.1075 0.06253 0.05922 0.0088 0.0079 1.0000 13.750 1.0910 0.06853 0.06540 0.0059 0.0079 1.0000 14.000 1.0720 0.07536 0.07241 0.0022 0.0079 1.0000 14.250 1.0508 0.08333 0.08056 -0.0025 0.0080 1.0000 14.500 1.0269 0.09252 0.08991 -0.0082 0.0082 1.0000 14.750 0.9998 0.10309 0.10062 -0.0145 0.0084 1.0000