XFOIL Version 6.96 Calculated polar for: HQ 0/7 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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.000 -0.6682 0.08759 0.08286 0.0050 1.0000 0.0125 -8.750 -0.6738 0.08131 0.07664 0.0004 1.0000 0.0122 -8.500 -0.6837 0.07408 0.06944 -0.0060 1.0000 0.0119 -8.250 -0.6959 0.06802 0.06330 -0.0093 1.0000 0.0115 -8.000 -0.7029 0.06213 0.05726 -0.0114 1.0000 0.0112 -7.750 -0.7056 0.05664 0.05152 -0.0125 1.0000 0.0109 -7.500 -0.7040 0.05145 0.04603 -0.0128 1.0000 0.0107 -7.250 -0.6982 0.04662 0.04081 -0.0127 1.0000 0.0106 -7.000 -0.6889 0.04215 0.03590 -0.0122 1.0000 0.0105 -6.750 -0.6761 0.03805 0.03131 -0.0114 1.0000 0.0105 -6.500 -0.6605 0.03426 0.02701 -0.0106 1.0000 0.0107 -6.250 -0.6424 0.03090 0.02316 -0.0097 1.0000 0.0111 -6.000 -0.6217 0.02803 0.01986 -0.0088 1.0000 0.0117 -5.750 -0.5993 0.02554 0.01699 -0.0079 1.0000 0.0127 -5.500 -0.5755 0.02390 0.01505 -0.0071 1.0000 0.0165 -5.250 -0.5521 0.02221 0.01312 -0.0064 1.0000 0.0227 -5.000 -0.5282 0.02092 0.01163 -0.0057 1.0000 0.0333 -4.750 -0.5058 0.01917 0.00982 -0.0046 1.0000 0.0392 -4.500 -0.4829 0.01790 0.00835 -0.0035 1.0000 0.0448 -4.250 -0.4598 0.01695 0.00731 -0.0028 1.0000 0.0563 -4.000 -0.4371 0.01592 0.00633 -0.0020 1.0000 0.0739 -3.750 -0.4147 0.01481 0.00528 -0.0012 1.0000 0.1159 -3.500 -0.3948 0.01346 0.00465 -0.0004 1.0000 0.2605 -3.250 -0.3783 0.01205 0.00433 0.0013 1.0000 0.4791 -3.000 -0.3602 0.01148 0.00433 0.0038 1.0000 0.6284 -2.750 -0.3401 0.01127 0.00428 0.0061 1.0000 0.7075 -2.500 -0.3203 0.01114 0.00425 0.0086 1.0000 0.7659 -2.250 -0.2995 0.01104 0.00415 0.0106 1.0000 0.8046 -2.000 -0.2765 0.01094 0.00398 0.0119 1.0000 0.8282 -1.750 -0.2530 0.01084 0.00381 0.0130 1.0000 0.8500 -1.500 -0.2283 0.01077 0.00360 0.0139 1.0000 0.8747 -1.250 -0.1984 0.01075 0.00356 0.0139 1.0000 0.9035 -1.000 -0.1614 0.01076 0.00353 0.0122 1.0000 0.9309 -0.750 -0.1216 0.01075 0.00347 0.0095 1.0000 0.9512 -0.500 -0.0814 0.01074 0.00340 0.0066 1.0000 0.9680 -0.250 -0.0401 0.01071 0.00335 0.0033 1.0000 0.9827 0.000 0.0000 0.01069 0.00332 0.0000 1.0000 1.0000 0.250 0.0401 0.01070 0.00334 -0.0033 0.9827 1.0000 0.500 0.0814 0.01073 0.00340 -0.0066 0.9680 1.0000 0.750 0.1215 0.01075 0.00347 -0.0095 0.9511 1.0000 1.000 0.1614 0.01075 0.00352 -0.0122 0.9308 1.0000 1.250 0.1982 0.01075 0.00356 -0.0138 0.9035 1.0000 1.500 0.2279 0.01077 0.00360 -0.0139 0.8747 1.0000 1.750 0.2526 0.01084 0.00380 -0.0129 0.8499 1.0000 2.000 0.2761 0.01094 0.00397 -0.0119 0.8282 1.0000 2.250 0.2990 0.01104 0.00415 -0.0106 0.8047 1.0000 2.500 0.3199 0.01114 0.00424 -0.0085 0.7663 1.0000 2.750 0.3396 0.01127 0.00428 -0.0060 0.7078 1.0000 3.000 0.3598 0.01148 0.00433 -0.0037 0.6298 1.0000 3.250 0.3779 0.01205 0.00433 -0.0012 0.4796 1.0000 3.500 0.3943 0.01348 0.00466 0.0005 0.2580 1.0000 3.750 0.4145 0.01480 0.00527 0.0012 0.1165 1.0000 4.000 0.4369 0.01592 0.00633 0.0020 0.0740 1.0000 4.250 0.4598 0.01695 0.00731 0.0028 0.0564 1.0000 4.500 0.4829 0.01791 0.00835 0.0035 0.0447 1.0000 4.750 0.5058 0.01919 0.00984 0.0046 0.0391 1.0000 5.000 0.5285 0.02090 0.01162 0.0057 0.0336 1.0000 5.250 0.5523 0.02222 0.01313 0.0063 0.0227 1.0000 5.500 0.5757 0.02391 0.01505 0.0071 0.0165 1.0000 5.750 0.5996 0.02557 0.01701 0.0078 0.0127 1.0000 6.000 0.6220 0.02805 0.01987 0.0087 0.0116 1.0000 6.250 0.6427 0.03092 0.02318 0.0096 0.0110 1.0000 6.500 0.6611 0.03429 0.02705 0.0105 0.0107 1.0000 6.750 0.6766 0.03809 0.03136 0.0113 0.0105 1.0000 7.000 0.6892 0.04225 0.03600 0.0120 0.0105 1.0000 7.250 0.6989 0.04669 0.04089 0.0125 0.0106 1.0000 7.500 0.7045 0.05152 0.04610 0.0126 0.0107 1.0000 7.750 0.7064 0.05678 0.05168 0.0122 0.0110 1.0000 8.000 0.7039 0.06223 0.05736 0.0112 0.0112 1.0000 8.250 0.6967 0.06818 0.06347 0.0090 0.0115 1.0000 8.500 0.6844 0.07427 0.06963 0.0056 0.0119 1.0000 8.750 0.6750 0.08136 0.07668 -0.0007 0.0121 1.0000 9.000 0.6695 0.08765 0.08292 -0.0053 0.0124 1.0000