XFOIL Version 6.96 Calculated polar for: GIII BL126 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.000 -0.5637 0.09466 0.08826 0.0096 1.0000 0.3044 -7.750 -0.5601 0.09100 0.08466 0.0106 1.0000 0.3251 -7.500 -0.5705 0.08856 0.08232 0.0110 1.0000 0.3495 -7.250 -0.5448 0.08428 0.07798 0.0148 1.0000 0.3829 -5.750 -0.5397 0.04897 0.04114 -0.0208 1.0000 0.1476 -5.500 -0.5229 0.04444 0.03607 -0.0199 1.0000 0.1310 -5.250 -0.5059 0.04088 0.03171 -0.0182 1.0000 0.1227 -5.000 -0.4889 0.03811 0.02836 -0.0164 1.0000 0.1237 -4.750 -0.4703 0.03497 0.02511 -0.0151 1.0000 0.1271 -4.500 -0.4494 0.03228 0.02201 -0.0135 1.0000 0.1280 -4.250 -0.4273 0.02996 0.01933 -0.0120 1.0000 0.1329 -4.000 -0.4060 0.02803 0.01718 -0.0106 1.0000 0.1461 -3.750 -0.3813 0.02597 0.01492 -0.0094 1.0000 0.1599 -3.500 -0.3564 0.02420 0.01315 -0.0083 1.0000 0.1897 -3.250 -0.1262 0.01893 0.00991 -0.0315 1.0000 1.0000 -3.000 -0.1071 0.01861 0.00927 -0.0309 1.0000 1.0000 -2.750 -0.0891 0.01834 0.00877 -0.0301 1.0000 1.0000 -2.500 -0.0726 0.01812 0.00836 -0.0290 1.0000 1.0000 -2.250 -0.0586 0.01795 0.00806 -0.0275 1.0000 1.0000 -2.000 -0.0487 0.01785 0.00785 -0.0254 1.0000 1.0000 -1.750 -0.0442 0.01783 0.00775 -0.0223 1.0000 1.0000 -1.500 -0.0440 0.01786 0.00768 -0.0185 1.0000 1.0000 -1.250 -0.0459 0.01793 0.00767 -0.0144 1.0000 1.0000 -1.000 -0.0473 0.01800 0.00765 -0.0104 1.0000 1.0000 -0.750 -0.0465 0.01808 0.00763 -0.0067 1.0000 1.0000 -0.500 -0.0429 0.01819 0.00763 -0.0035 1.0000 1.0000 -0.250 -0.0364 0.01832 0.00765 -0.0008 1.0000 1.0000 0.000 -0.0266 0.01850 0.00771 0.0013 1.0000 1.0000 0.250 -0.0142 0.01872 0.00781 0.0029 1.0000 1.0000 0.500 0.0001 0.01898 0.00798 0.0042 1.0000 1.0000 0.750 0.0158 0.01928 0.00821 0.0052 1.0000 1.0000 1.000 0.0325 0.01963 0.00849 0.0060 1.0000 1.0000 1.250 0.0498 0.02001 0.00883 0.0066 1.0000 1.0000 1.500 0.0676 0.02044 0.00922 0.0071 1.0000 1.0000 1.750 0.0857 0.02091 0.00968 0.0074 1.0000 1.0000 2.000 0.1041 0.02142 0.01019 0.0077 1.0000 1.0000 2.250 0.1225 0.02198 0.01077 0.0078 1.0000 1.0000 2.500 0.1409 0.02260 0.01142 0.0079 1.0000 1.0000 2.750 0.1592 0.02327 0.01216 0.0078 1.0000 1.0000 3.000 0.2189 0.02463 0.01369 -0.0002 0.9829 1.0000 3.250 0.2731 0.02578 0.01505 -0.0068 0.9618 1.0000 3.500 0.3334 0.02692 0.01651 -0.0141 0.9391 1.0000 3.750 0.3916 0.02782 0.01774 -0.0205 0.9135 1.0000 4.000 0.4493 0.02846 0.01883 -0.0260 0.8851 1.0000 4.250 0.5057 0.02880 0.01963 -0.0304 0.8528 1.0000 4.500 0.5715 0.02832 0.01976 -0.0341 0.8158 1.0000 4.750 0.6234 0.02664 0.01865 -0.0323 0.7679 1.0000 5.000 0.6563 0.02407 0.01643 -0.0256 0.7088 1.0000 5.250 0.6755 0.02190 0.01445 -0.0180 0.6382 1.0000 5.500 0.6813 0.02071 0.01257 -0.0084 0.4538 1.0000 5.750 0.6750 0.02443 0.01404 -0.0020 0.2342 1.0000 6.000 0.6913 0.02696 0.01597 0.0002 0.1697 1.0000 6.250 0.7194 0.02937 0.01820 0.0013 0.1419 1.0000 6.500 0.7460 0.03163 0.02049 0.0022 0.1231 1.0000 6.750 0.7744 0.03459 0.02363 0.0029 0.1155 1.0000 7.000 0.7985 0.03745 0.02689 0.0041 0.1095 1.0000 7.250 0.8191 0.04068 0.03033 0.0051 0.1033 1.0000 7.500 0.8364 0.04418 0.03443 0.0067 0.1019 1.0000 7.750 0.8515 0.04821 0.03901 0.0082 0.1026 1.0000 8.000 0.8647 0.05266 0.04384 0.0095 0.1040 1.0000 8.250 0.8656 0.05717 0.04919 0.0114 0.1086 1.0000 8.500 0.8617 0.06258 0.05513 0.0124 0.1136 1.0000 8.750 0.8648 0.06772 0.06047 0.0129 0.1167 1.0000 9.000 0.8382 0.07411 0.06736 0.0121 0.1264 1.0000 9.500 0.7987 0.08755 0.08098 0.0070 0.1474 1.0000 9.750 0.6877 0.08079 0.07444 0.0157 0.1334 1.0000 10.000 0.6379 0.09150 0.08504 0.0074 0.1412 1.0000