XFOIL Version 6.96 Calculated polar for: HQ 0/9 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 ------ -------- --------- --------- -------- -------- -------- -9.750 -0.5705 0.10293 0.09855 0.0068 1.0000 0.1215 -9.500 -0.5902 0.09764 0.09333 0.0021 1.0000 0.1224 -9.250 -0.6040 0.09160 0.08736 -0.0015 1.0000 0.1235 -9.000 -0.5702 0.08900 0.08471 0.0044 1.0000 0.1401 -8.750 -0.5786 0.08439 0.08013 0.0022 1.0000 0.1473 -8.500 -0.6119 0.07837 0.07422 -0.0044 1.0000 0.1489 -8.250 -0.5885 0.07463 0.07042 -0.0004 1.0000 0.1594 -7.250 -0.7210 0.04714 0.04064 -0.0147 1.0000 0.0562 -7.000 -0.7085 0.04240 0.03552 -0.0138 1.0000 0.0558 -6.750 -0.6954 0.03848 0.03094 -0.0125 1.0000 0.0576 -6.500 -0.6772 0.03427 0.02624 -0.0113 1.0000 0.0578 -6.250 -0.6566 0.03053 0.02220 -0.0104 1.0000 0.0599 -6.000 -0.6340 0.02857 0.01997 -0.0095 1.0000 0.0671 -5.750 -0.6090 0.02609 0.01695 -0.0083 1.0000 0.0690 -5.500 -0.5837 0.02334 0.01401 -0.0075 1.0000 0.0726 -5.250 -0.5590 0.02206 0.01256 -0.0068 1.0000 0.0819 -5.000 -0.5346 0.02007 0.01067 -0.0059 1.0000 0.0893 -4.750 -0.5115 0.01849 0.00913 -0.0048 1.0000 0.0998 -4.500 -0.4905 0.01707 0.00784 -0.0034 1.0000 0.1222 -4.250 -0.4735 0.01517 0.00664 -0.0018 1.0000 0.1995 -4.000 -0.4668 0.01267 0.00596 0.0016 1.0000 0.4994 -3.750 -0.4522 0.01235 0.00616 0.0055 1.0000 0.6488 -3.500 -0.4353 0.01237 0.00622 0.0090 1.0000 0.7185 -3.250 -0.4190 0.01246 0.00633 0.0127 1.0000 0.7671 -3.000 -0.4033 0.01253 0.00637 0.0165 1.0000 0.8057 -2.750 -0.3886 0.01263 0.00646 0.0207 1.0000 0.8421 -2.500 -0.3702 0.01280 0.00659 0.0245 1.0000 0.8776 -2.250 -0.3390 0.01295 0.00657 0.0251 1.0000 0.9048 -2.000 -0.2994 0.01299 0.00644 0.0232 1.0000 0.9240 -1.750 -0.2539 0.01302 0.00630 0.0198 1.0000 0.9409 -1.500 -0.1933 0.01307 0.00614 0.0135 1.0000 0.9548 -1.250 -0.1138 0.01306 0.00593 0.0036 1.0000 0.9706 -1.000 -0.0506 0.01280 0.00557 -0.0038 1.0000 0.9848 -0.750 0.0025 0.01245 0.00517 -0.0097 1.0000 0.9964 -0.500 0.0257 0.01222 0.00495 -0.0104 1.0000 1.0000 -0.250 0.0238 0.01213 0.00493 -0.0068 1.0000 1.0000 0.000 0.0000 0.01215 0.00498 0.0000 1.0000 1.0000 0.250 -0.0238 0.01213 0.00493 0.0068 1.0000 1.0000 0.500 -0.0257 0.01221 0.00495 0.0104 1.0000 1.0000 0.750 -0.0025 0.01245 0.00516 0.0097 0.9965 1.0000 1.000 0.0506 0.01280 0.00557 0.0038 0.9848 1.0000 1.250 0.1138 0.01306 0.00593 -0.0037 0.9706 1.0000 1.500 0.1934 0.01307 0.00613 -0.0135 0.9548 1.0000 1.750 0.2539 0.01301 0.00629 -0.0198 0.9409 1.0000 2.000 0.2993 0.01299 0.00644 -0.0232 0.9240 1.0000 2.250 0.3389 0.01295 0.00656 -0.0251 0.9048 1.0000 2.500 0.3699 0.01280 0.00659 -0.0244 0.8776 1.0000 2.750 0.3883 0.01263 0.00646 -0.0206 0.8421 1.0000 3.000 0.4030 0.01253 0.00637 -0.0164 0.8058 1.0000 3.250 0.4187 0.01246 0.00633 -0.0127 0.7672 1.0000 3.500 0.4351 0.01237 0.00622 -0.0089 0.7190 1.0000 3.750 0.4519 0.01235 0.00616 -0.0054 0.6491 1.0000 4.000 0.4666 0.01267 0.00596 -0.0016 0.5001 1.0000 4.250 0.4734 0.01517 0.00664 0.0018 0.1996 1.0000 4.500 0.4904 0.01708 0.00784 0.0035 0.1224 1.0000 4.750 0.5115 0.01848 0.00912 0.0048 0.0999 1.0000 5.000 0.5346 0.02007 0.01067 0.0059 0.0894 1.0000 5.250 0.5590 0.02206 0.01255 0.0068 0.0820 1.0000 5.500 0.5837 0.02334 0.01401 0.0075 0.0726 1.0000 5.750 0.6090 0.02609 0.01696 0.0083 0.0690 1.0000 6.000 0.6341 0.02858 0.01997 0.0095 0.0670 1.0000 6.250 0.6568 0.03055 0.02223 0.0104 0.0600 1.0000 6.500 0.6773 0.03430 0.02626 0.0112 0.0578 1.0000 6.750 0.6961 0.03824 0.03073 0.0125 0.0578 1.0000 7.000 0.7087 0.04257 0.03565 0.0137 0.0560 1.0000 7.250 0.7212 0.04720 0.04071 0.0147 0.0562 1.0000 8.500 0.7293 0.08324 0.07868 0.0103 0.1227 1.0000 8.750 0.6939 0.08757 0.08312 0.0044 0.1219 1.0000 9.000 0.6646 0.09368 0.08910 -0.0047 0.1195 1.0000