XFOIL Version 6.96 Calculated polar for: HAM-STD HS1-404 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 ------ -------- --------- --------- -------- -------- -------- -8.000 -0.4470 0.09829 0.09601 -0.0089 1.0000 0.0077 -7.750 -0.4402 0.09530 0.09304 -0.0107 1.0000 0.0078 -7.500 -0.4342 0.09226 0.09003 -0.0124 1.0000 0.0079 -5.250 -0.3305 0.05537 0.05315 -0.0416 0.9998 0.0088 -5.000 -0.2940 0.05010 0.04779 -0.0492 0.9977 0.0097 -4.750 -0.2551 0.04547 0.04306 -0.0561 0.9958 0.0107 -4.500 -0.2124 0.04031 0.03774 -0.0632 0.9942 0.0120 -4.250 -0.1695 0.03489 0.03211 -0.0695 0.9918 0.0136 -4.000 -0.1252 0.02906 0.02596 -0.0748 0.9897 0.0156 -3.750 -0.0820 0.02607 0.02260 -0.0774 0.9877 0.0183 -3.500 -0.0428 0.01890 0.01480 -0.0826 0.9864 0.0217 -3.250 -0.0026 0.01493 0.01020 -0.0843 0.9852 0.0147 -2.750 0.0711 0.01175 0.00641 -0.0877 0.9828 0.0167 -2.500 0.1058 0.01017 0.00455 -0.0888 0.9805 0.0149 -2.250 0.1403 0.00914 0.00337 -0.0900 0.9776 0.0138 -2.000 0.1753 0.00849 0.00259 -0.0913 0.9746 0.0135 -1.750 0.2111 0.00808 0.00204 -0.0927 0.9715 0.0147 -1.500 0.2443 0.00736 0.00174 -0.0939 0.9659 0.1498 -1.250 0.2754 0.00714 0.00172 -0.0946 0.9578 0.2222 -1.000 0.3035 0.00704 0.00151 -0.0944 0.9476 0.2280 -0.750 0.3307 0.00683 0.00147 -0.0942 0.9381 0.2809 -0.500 0.3573 0.00655 0.00140 -0.0939 0.9284 0.3568 -0.250 0.3762 0.00499 0.00140 -0.0918 0.9182 1.0000 0.000 0.4031 0.00503 0.00137 -0.0914 0.9093 1.0000 0.250 0.4300 0.00509 0.00135 -0.0910 0.9015 1.0000 0.500 0.4570 0.00514 0.00138 -0.0907 0.8930 1.0000 0.750 0.4840 0.00519 0.00140 -0.0904 0.8841 1.0000 1.000 0.5108 0.00526 0.00142 -0.0900 0.8753 1.0000 1.250 0.5376 0.00532 0.00146 -0.0897 0.8653 1.0000 1.500 0.5642 0.00538 0.00150 -0.0892 0.8525 1.0000 1.750 0.5898 0.00545 0.00151 -0.0884 0.8308 1.0000 2.000 0.6155 0.00553 0.00154 -0.0877 0.8058 1.0000 2.250 0.6408 0.00566 0.00161 -0.0868 0.7699 1.0000 2.500 0.6666 0.00581 0.00166 -0.0862 0.7278 1.0000 2.750 0.6911 0.00614 0.00173 -0.0853 0.6502 1.0000 3.000 0.7123 0.00711 0.00201 -0.0841 0.4968 1.0000 3.250 0.7336 0.00831 0.00244 -0.0834 0.3270 1.0000 3.500 0.7557 0.00948 0.00291 -0.0829 0.1703 1.0000 3.750 0.7780 0.01071 0.00358 -0.0823 0.0564 1.0000 4.000 0.8031 0.01140 0.00432 -0.0817 0.0380 1.0000 4.250 0.8285 0.01201 0.00500 -0.0812 0.0293 1.0000 4.500 0.8528 0.01285 0.00592 -0.0805 0.0246 1.0000 4.750 0.8800 0.01281 0.00587 -0.0806 0.0179 1.0000 5.000 0.9055 0.01330 0.00633 -0.0802 0.0080 1.0000 5.250 0.9252 0.01557 0.00886 -0.0782 0.0045 1.0000 5.500 0.9470 0.01753 0.01108 -0.0766 0.0040 1.0000 5.750 0.9687 0.02039 0.01429 -0.0749 0.0039 1.0000 6.000 0.9895 0.02430 0.01869 -0.0728 0.0041 1.0000 6.250 1.0054 0.02979 0.02481 -0.0701 0.0044 1.0000 6.500 0.9845 0.02081 0.01642 -0.0623 0.0053 1.0000 6.750 1.0094 0.04948 0.04581 -0.0623 0.0118 1.0000 7.000 1.0163 0.05415 0.05077 -0.0604 0.0118 1.0000 7.250 1.0207 0.05882 0.05571 -0.0585 0.0117 1.0000 7.500 1.0223 0.06349 0.06062 -0.0569 0.0117 1.0000 7.750 1.0210 0.06813 0.06548 -0.0554 0.0117 1.0000 8.000 1.0164 0.07275 0.07028 -0.0542 0.0117 1.0000 8.250 1.0083 0.07731 0.07499 -0.0534 0.0116 1.0000 8.500 0.9935 0.08137 0.07917 -0.0523 0.0116 1.0000 8.750 0.9779 0.08618 0.08408 -0.0533 0.0116 1.0000