XFOIL Version 6.96 Calculated polar for: FX 61-147 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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -10.750 -0.6120 0.04387 0.04100 -0.0772 0.9947 0.0039 -10.500 -0.6310 0.03645 0.03309 -0.0826 0.9869 0.0039 -10.250 -0.6362 0.03143 0.02763 -0.0851 0.9769 0.0039 -10.000 -0.6263 0.02783 0.02363 -0.0866 0.9684 0.0039 -9.500 -0.5892 0.02329 0.01850 -0.0882 0.9555 0.0040 -9.250 -0.5645 0.02159 0.01658 -0.0892 0.9505 0.0041 -9.000 -0.5417 0.02025 0.01504 -0.0895 0.9438 0.0042 -8.750 -0.5141 0.01885 0.01349 -0.0907 0.9383 0.0044 -8.500 -0.4822 0.01789 0.01244 -0.0925 0.9336 0.0047 -8.250 -0.4511 0.01708 0.01152 -0.0939 0.9261 0.0049 -8.000 -0.4122 0.01620 0.01054 -0.0970 0.9209 0.0053 -7.750 -0.3753 0.01537 0.00960 -0.0996 0.9126 0.0058 -7.500 -0.3301 0.01451 0.00862 -0.1039 0.9057 0.0065 -7.250 -0.2862 0.01393 0.00798 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2.750 0.8579 0.01025 0.00354 -0.1233 0.4407 0.6917 3.000 0.8852 0.01036 0.00367 -0.1233 0.4351 0.6940 3.250 0.9121 0.01051 0.00380 -0.1232 0.4297 0.6965 3.500 0.9398 0.01062 0.00392 -0.1233 0.4239 0.6989 3.750 0.9669 0.01076 0.00405 -0.1233 0.4183 0.7013 4.000 0.9936 0.01089 0.00421 -0.1232 0.4135 0.7031 4.250 1.0205 0.01101 0.00437 -0.1232 0.4081 0.7050 4.500 1.0468 0.01117 0.00454 -0.1230 0.4025 0.7069 4.750 1.0735 0.01131 0.00472 -0.1229 0.3969 0.7090 5.000 1.1001 0.01146 0.00490 -0.1228 0.3909 0.7113 5.250 1.1260 0.01164 0.00509 -0.1226 0.3857 0.7136 5.500 1.1528 0.01179 0.00529 -0.1225 0.3791 0.7157 5.750 1.1778 0.01199 0.00550 -0.1222 0.3705 0.7175 6.000 1.2026 0.01219 0.00571 -0.1218 0.3592 0.7193 6.250 1.2269 0.01243 0.00595 -0.1213 0.3463 0.7211 6.500 1.2511 0.01268 0.00622 -0.1208 0.3346 0.7231 6.750 1.2747 0.01296 0.00649 -0.1202 0.3218 0.7253 7.000 1.2956 0.01340 0.00683 -0.1192 0.2977 0.7274 7.250 1.3145 0.01397 0.00725 -0.1179 0.2662 0.7296 7.500 1.3327 0.01459 0.00772 -0.1165 0.2362 0.7315 7.750 1.3498 0.01524 0.00827 -0.1149 0.2087 0.7333 8.000 1.3607 0.01623 0.00902 -0.1124 0.1648 0.7350 8.250 1.3665 0.01734 0.00990 -0.1089 0.1237 0.7368 8.500 1.3724 0.01845 0.01082 -0.1056 0.0915 0.7390 8.750 1.3809 0.01941 0.01168 -0.1027 0.0682 0.7412 9.000 1.3905 0.02030 0.01251 -0.1001 0.0538 0.7434 9.250 1.4005 0.02117 0.01335 -0.0976 0.0420 0.7454 9.500 1.4089 0.02212 0.01431 -0.0950 0.0317 0.7471 9.750 1.4174 0.02308 0.01528 -0.0926 0.0249 0.7489 10.000 1.4248 0.02414 0.01637 -0.0901 0.0194 0.7509 10.250 1.4336 0.02517 0.01745 -0.0879 0.0156 0.7532 10.500 1.4400 0.02639 0.01872 -0.0857 0.0116 0.7554 10.750 1.4431 0.02791 0.02025 -0.0833 0.0065 0.7575 11.000 1.4450 0.02961 0.02200 -0.0811 0.0042 0.7595 11.250 1.4483 0.03132 0.02379 -0.0792 0.0035 0.7613 11.500 1.4523 0.03307 0.02565 -0.0777 0.0031 0.7632 11.750 1.4558 0.03499 0.02768 -0.0763 0.0029 0.7653 12.000 1.4582 0.03712 0.02994 -0.0752 0.0027 0.7677 12.250 1.4593 0.03950 0.03244 -0.0742 0.0025 0.7700 12.500 1.4600 0.04207 0.03513 -0.0735 0.0024 0.7723 12.750 1.4597 0.04489 0.03806 -0.0730 0.0023 0.7744 13.000 1.4582 0.04796 0.04127 -0.0728 0.0022 0.7762 13.250 1.4577 0.05106 0.04449 -0.0727 0.0021 0.7781 13.500 1.4569 0.05429 0.04785 -0.0729 0.0021 0.7801 13.750 1.4554 0.05772 0.05140 -0.0732 0.0021 0.7823 14.000 1.4533 0.06135 0.05516 -0.0736 0.0020 0.7846 14.250 1.4505 0.06518 0.05911 -0.0743 0.0020 0.7870 14.500 1.4475 0.06915 0.06321 -0.0751 0.0020 0.7892 14.750 1.4438 0.07331 0.06750 -0.0761 0.0019 0.7913 15.000 1.4397 0.07764 0.07196 -0.0772 0.0019 0.7934 15.250 1.4354 0.08212 0.07657 -0.0785 0.0018 0.7955 15.500 1.4307 0.08676 0.08136 -0.0800 0.0018 0.7977 15.750 1.4257 0.09155 0.08628 -0.0816 0.0018 0.7999 16.000 1.4203 0.09651 0.09136 -0.0834 0.0017 0.8021 16.500 1.4102 0.10654 0.10165 -0.0873 0.0017 0.8066 16.750 1.4047 0.11174 0.10698 -0.0895 0.0016 0.8090 17.000 1.3996 0.11697 0.11233 -0.0918 0.0016 0.8117 17.250 1.3942 0.12231 0.11780 -0.0942 0.0016 0.8148 17.500 1.3889 0.12764 0.12325 -0.0968 0.0016 0.8179