XFOIL Version 6.96 Calculated polar for: BOEING HSNLF 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 ------ -------- --------- --------- -------- -------- -------- -9.500 -0.5568 0.07979 0.07822 -0.0431 1.0014 0.0082 -9.250 -0.5775 0.07503 0.07346 -0.0439 1.0014 0.0082 -9.000 -0.5914 0.06932 0.06769 -0.0481 1.0003 0.0082 -8.750 -0.5985 0.06347 0.06172 -0.0544 0.9975 0.0082 -8.500 -0.5948 0.05796 0.05604 -0.0586 0.9949 0.0082 -8.250 -0.5895 0.05320 0.05109 -0.0607 0.9917 0.0082 -8.000 -0.5795 0.04853 0.04620 -0.0626 0.9894 0.0083 -7.750 -0.5655 0.04405 0.04147 -0.0643 0.9878 0.0083 -7.500 -0.5480 0.03983 0.03696 -0.0660 0.9866 0.0083 -5.500 -0.3652 0.01312 0.00787 -0.0678 0.9741 0.0054 -5.250 -0.3345 0.01257 0.00726 -0.0687 0.9734 0.0052 -5.000 -0.3079 0.01131 0.00590 -0.0689 0.9719 0.0050 -4.750 -0.2865 0.01052 0.00501 -0.0679 0.9682 0.0049 -4.500 -0.2590 0.00989 0.00428 -0.0681 0.9660 0.0049 -4.250 -0.2300 0.00942 0.00372 -0.0685 0.9643 0.0049 -4.000 -0.2004 0.00907 0.00331 -0.0691 0.9628 0.0051 -3.750 -0.1704 0.00877 0.00295 -0.0698 0.9615 0.0068 -3.500 -0.1408 0.00846 0.00270 -0.0704 0.9603 0.0236 -3.250 -0.1147 0.00804 0.00248 -0.0705 0.9579 0.0810 -3.000 -0.0896 0.00743 0.00224 -0.0705 0.9550 0.1925 -2.750 -0.0650 0.00623 0.00188 -0.0709 0.9525 0.4289 -2.500 -0.0378 0.00567 0.00179 -0.0713 0.9504 0.5728 -2.250 -0.0087 0.00556 0.00176 -0.0716 0.9485 0.6087 -2.000 0.0194 0.00557 0.00185 -0.0717 0.9463 0.6453 -1.750 0.0462 0.00560 0.00190 -0.0715 0.9429 0.6582 -1.500 0.0742 0.00559 0.00189 -0.0716 0.9401 0.6659 -1.250 0.1029 0.00556 0.00185 -0.0718 0.9377 0.6699 -1.000 0.1319 0.00554 0.00181 -0.0722 0.9355 0.6735 -0.750 0.1607 0.00555 0.00180 -0.0725 0.9333 0.6767 -0.500 0.1878 0.00550 0.00179 -0.0724 0.9295 0.6800 -0.250 0.2155 0.00545 0.00175 -0.0724 0.9250 0.6831 0.000 0.2438 0.00538 0.00167 -0.0725 0.9207 0.6864 0.250 0.2709 0.00535 0.00164 -0.0723 0.9147 0.6897 0.500 0.2987 0.00526 0.00156 -0.0723 0.9088 0.6928 0.750 0.3251 0.00514 0.00147 -0.0718 0.8978 0.6959 1.000 0.3514 0.00503 0.00132 -0.0713 0.8800 0.6991 1.250 0.3788 0.00502 0.00130 -0.0712 0.8670 0.7025 1.500 0.4065 0.00501 0.00130 -0.0712 0.8544 0.7058 1.750 0.4327 0.00503 0.00127 -0.0707 0.8259 0.7089 2.000 0.4576 0.00518 0.00130 -0.0700 0.7813 0.7122 2.250 0.4709 0.00606 0.00146 -0.0669 0.6108 0.7159 2.500 0.4852 0.00716 0.00179 -0.0645 0.4197 0.7193 2.750 0.5042 0.00798 0.00208 -0.0632 0.2796 0.7229 3.000 0.5249 0.00872 0.00237 -0.0621 0.1622 0.7265 3.250 0.5484 0.00922 0.00262 -0.0615 0.0975 0.7303 3.500 0.5728 0.00960 0.00284 -0.0611 0.0577 0.7340 3.750 0.5981 0.00989 0.00307 -0.0608 0.0365 0.7381 4.000 0.6224 0.01032 0.00338 -0.0602 0.0120 0.7425 4.250 0.6477 0.01065 0.00372 -0.0598 0.0080 0.7467 4.500 0.6731 0.01094 0.00410 -0.0594 0.0074 0.7512 4.750 0.6983 0.01127 0.00450 -0.0590 0.0071 0.7562 5.000 0.7232 0.01164 0.00495 -0.0585 0.0067 0.7615 5.250 0.7475 0.01206 0.00548 -0.0578 0.0062 0.7673 5.500 0.7715 0.01255 0.00604 -0.0572 0.0057 0.7737 5.750 0.7944 0.01317 0.00677 -0.0563 0.0053 0.7807 6.250 0.8367 0.01512 0.00903 -0.0537 0.0049 0.7984 6.500 0.8587 0.01614 0.01022 -0.0526 0.0050 0.8099 6.750 0.8808 0.01733 0.01162 -0.0514 0.0051 0.8243 7.250 0.8901 0.03422 0.03043 -0.0401 0.0116 0.8601 7.500 0.8637 0.02349 0.02008 -0.0318 0.0108 0.8419 7.750 0.8714 0.02665 0.02359 -0.0286 0.0102 0.8668 8.000 0.8729 0.02997 0.02727 -0.0247 0.0097 0.9108 8.250 0.8734 0.03369 0.03129 -0.0211 0.0093 0.9986 8.500 0.8717 0.03767 0.03545 -0.0179 0.0091 0.9986 8.750 0.8663 0.04165 0.03961 -0.0146 0.0089 0.9986 9.000 0.8571 0.04550 0.04360 -0.0111 0.0088 0.9986 9.250 0.8394 0.04899 0.04722 -0.0067 0.0087 0.9986 9.500 0.8170 0.05267 0.05102 -0.0027 0.0087 0.9986 9.750 0.7898 0.05721 0.05567 0.0001 0.0087 0.9986 10.000 0.7593 0.06271 0.06129 0.0011 0.0088 0.9986 10.250 0.7226 0.07046 0.06914 -0.0007 0.0091 0.9986 10.500 0.6803 0.08393 0.08268 -0.0099 0.0096 0.9986