XFOIL Version 6.96 Calculated polar for: GOE 100 (SOPWITH) 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 ------ -------- --------- --------- -------- -------- -------- -7.500 -0.4050 0.08586 0.08440 -0.0195 1.0000 0.0076 -7.250 -0.4072 0.08284 0.08141 -0.0202 1.0000 0.0076 -7.000 -0.4111 0.08000 0.07860 -0.0202 1.0000 0.0076 -6.750 -0.3919 0.07463 0.07321 -0.0269 0.9983 0.0076 -6.500 -0.3647 0.06829 0.06683 -0.0356 0.9960 0.0076 -6.250 -0.3412 0.05895 0.05741 -0.0465 0.9932 0.0084 -6.000 -0.3141 0.05553 0.05393 -0.0511 0.9903 0.0089 -5.750 -0.2836 0.05134 0.04966 -0.0564 0.9876 0.0096 -5.500 -0.2497 0.04635 0.04456 -0.0622 0.9854 0.0106 -5.250 -0.2105 0.04306 0.04116 -0.0662 0.9840 0.0128 -4.750 -0.1513 0.01547 0.01181 -0.0749 0.9740 0.0083 -4.500 -0.1189 0.01410 0.01016 -0.0758 0.9708 0.0079 -4.250 -0.0904 0.01253 0.00830 -0.0758 0.9642 0.0078 -4.000 -0.0619 0.00995 0.00525 -0.0758 0.9579 0.0084 -3.750 -0.0341 0.00913 0.00431 -0.0757 0.9476 0.0097 -3.500 -0.0064 0.00864 0.00370 -0.0754 0.9347 0.0106 -3.250 0.0206 0.00820 0.00313 -0.0749 0.9214 0.0114 -3.000 0.0478 0.00791 0.00276 -0.0746 0.9110 0.0125 -2.750 0.0751 0.00752 0.00221 -0.0743 0.9021 0.0144 -2.500 0.1025 0.00720 0.00182 -0.0739 0.8923 0.0190 -2.250 0.1297 0.00701 0.00170 -0.0737 0.8817 0.0408 -2.000 0.1570 0.00699 0.00164 -0.0735 0.8698 0.0488 -1.750 0.1845 0.00696 0.00155 -0.0733 0.8594 0.0535 -1.500 0.2122 0.00692 0.00148 -0.0732 0.8494 0.0556 -1.250 0.2398 0.00684 0.00134 -0.0730 0.8387 0.0588 -1.000 0.2674 0.00673 0.00123 -0.0729 0.8281 0.0654 -0.750 0.2950 0.00668 0.00112 -0.0728 0.8150 0.0685 -0.500 0.3223 0.00656 0.00103 -0.0726 0.7996 0.0899 -0.250 0.3494 0.00645 0.00102 -0.0725 0.7813 0.1513 0.000 0.3765 0.00645 0.00100 -0.0722 0.7566 0.1718 0.250 0.4031 0.00650 0.00097 -0.0719 0.7243 0.1901 0.500 0.4285 0.00662 0.00097 -0.0714 0.6676 0.2217 0.750 0.4522 0.00672 0.00107 -0.0709 0.5866 0.3589 1.000 0.4684 0.00561 0.00123 -0.0688 0.5310 0.8995 1.250 0.4972 0.00573 0.00129 -0.0688 0.4849 0.9890 1.500 0.5284 0.00600 0.00136 -0.0697 0.4412 1.0000 1.750 0.5541 0.00626 0.00144 -0.0694 0.4006 1.0000 2.000 0.5798 0.00653 0.00153 -0.0691 0.3642 1.0000 2.250 0.6054 0.00685 0.00164 -0.0687 0.3233 1.0000 2.500 0.6301 0.00727 0.00178 -0.0683 0.2682 1.0000 2.750 0.6560 0.00755 0.00192 -0.0680 0.2352 1.0000 3.000 0.6815 0.00789 0.00206 -0.0677 0.2003 1.0000 3.250 0.7065 0.00830 0.00223 -0.0674 0.1590 1.0000 3.750 0.7522 0.00983 0.00306 -0.0659 0.0160 1.0000 4.000 0.7789 0.01003 0.00331 -0.0657 0.0136 1.0000 4.250 0.8054 0.01027 0.00361 -0.0654 0.0124 1.0000 4.500 0.8316 0.01054 0.00393 -0.0651 0.0114 1.0000 4.750 0.8574 0.01089 0.00434 -0.0648 0.0105 1.0000 5.000 0.8826 0.01134 0.00486 -0.0643 0.0094 1.0000 5.250 0.9071 0.01193 0.00552 -0.0636 0.0085 1.0000 5.500 0.9311 0.01258 0.00625 -0.0628 0.0084 1.0000 5.750 0.9551 0.01324 0.00699 -0.0620 0.0088 1.0000 6.000 0.9766 0.01416 0.00797 -0.0610 0.0080 1.0000 6.250 0.9961 0.01536 0.00921 -0.0596 0.0077 1.0000 6.500 1.0151 0.01664 0.01054 -0.0581 0.0076 1.0000 6.750 1.0368 0.01756 0.01155 -0.0570 0.0066 1.0000 7.000 1.0593 0.01841 0.01246 -0.0560 0.0062 1.0000 11.000 1.1837 0.03990 0.03637 -0.0151 0.0045 1.0000 11.250 1.1696 0.04267 0.03927 -0.0118 0.0044 1.0000 11.500 1.1533 0.04580 0.04254 -0.0090 0.0044 1.0000 11.750 1.1338 0.04946 0.04635 -0.0068 0.0044 1.0000 12.000 1.1136 0.05364 0.05066 -0.0054 0.0044 1.0000 12.250 1.0900 0.05854 0.05572 -0.0049 0.0044 1.0000 12.500 1.0662 0.06400 0.06132 -0.0054 0.0044 1.0000 12.750 1.0404 0.07020 0.06767 -0.0070 0.0044 1.0000 13.000 1.0138 0.07715 0.07478 -0.0097 0.0045 1.0000 13.250 0.9878 0.08439 0.08215 -0.0130 0.0045 1.0000 13.500 0.9618 0.09211 0.09001 -0.0170 0.0046 1.0000 13.750 0.9344 0.10025 0.09829 -0.0216 0.0047 1.0000 14.000 0.9052 0.10867 0.10684 -0.0264 0.0049 1.0000 14.250 0.8725 0.11564 0.11392 -0.0301 0.0052 1.0000