XFOIL Version 6.96 Calculated polar for: GOE 445 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 ------ -------- --------- --------- -------- -------- -------- -8.750 -0.6363 0.08389 0.08171 -0.0042 1.0000 0.0054 -8.500 -0.6404 0.07815 0.07601 -0.0089 1.0000 0.0053 -8.000 -0.6562 0.06738 0.06518 -0.0163 1.0000 0.0053 -7.750 -0.6576 0.06229 0.06001 -0.0177 1.0000 0.0052 -7.500 -0.6571 0.05764 0.05520 -0.0181 1.0000 0.0052 -7.250 -0.6549 0.05354 0.05097 -0.0179 1.0000 0.0053 -7.000 -0.6502 0.04944 0.04669 -0.0173 1.0000 0.0054 -6.750 -0.6431 0.04562 0.04269 -0.0163 1.0000 0.0055 -6.500 -0.6340 0.04166 0.03850 -0.0149 1.0000 0.0055 -6.250 -0.6233 0.03781 0.03439 -0.0131 1.0000 0.0055 -5.750 -0.5975 0.03114 0.02719 -0.0090 1.0000 0.0041 -5.500 -0.5838 0.02728 0.02295 -0.0064 1.0000 0.0036 -5.000 -0.5475 0.02005 0.01482 -0.0006 1.0000 0.0028 -4.750 -0.5283 0.01764 0.01204 0.0015 1.0000 0.0028 -4.500 -0.5082 0.01535 0.00942 0.0035 1.0000 0.0029 -4.250 -0.4881 0.01376 0.00764 0.0051 1.0000 0.0031 -4.000 -0.4670 0.01280 0.00655 0.0065 1.0000 0.0034 -3.750 -0.4456 0.01206 0.00571 0.0079 1.0000 0.0041 -3.500 -0.4242 0.01141 0.00497 0.0092 1.0000 0.0052 -3.250 -0.3986 0.01049 0.00395 0.0096 0.9987 0.0057 -3.000 -0.3667 0.00987 0.00324 0.0086 0.9958 0.0069 -2.750 -0.3354 0.00943 0.00270 0.0078 0.9927 0.0089 -2.500 -0.3038 0.00898 0.00213 0.0070 0.9894 0.0153 -2.250 -0.2750 0.00806 0.00174 0.0062 0.9858 0.1516 -2.000 -0.2508 0.00676 0.00144 0.0060 0.9806 0.4169 -1.750 -0.2247 0.00600 0.00134 0.0060 0.9756 0.5948 -1.500 -0.1947 0.00580 0.00122 0.0054 0.9703 0.6313 -1.250 -0.1636 0.00559 0.00113 0.0047 0.9656 0.6716 -1.000 -0.1362 0.00529 0.00107 0.0048 0.9582 0.7469 -0.750 -0.1060 0.00510 0.00104 0.0044 0.9517 0.7961 -0.500 -0.0730 0.00496 0.00101 0.0033 0.9416 0.8333 -0.250 -0.0365 0.00484 0.00096 0.0017 0.9203 0.8678 0.000 0.0000 0.00480 0.00094 0.0000 0.8945 0.8944 0.250 0.0366 0.00484 0.00096 -0.0017 0.8676 0.9204 0.500 0.0729 0.00496 0.00101 -0.0033 0.8337 0.9414 0.750 0.1062 0.00510 0.00104 -0.0044 0.7959 0.9516 1.000 0.1362 0.00530 0.00107 -0.0048 0.7439 0.9582 1.250 0.1639 0.00559 0.00113 -0.0047 0.6726 0.9656 1.500 0.1950 0.00581 0.00122 -0.0055 0.6311 0.9704 1.750 0.2251 0.00601 0.00134 -0.0061 0.5948 0.9757 2.000 0.2515 0.00670 0.00145 -0.0061 0.4319 0.9805 2.250 0.2753 0.00807 0.00174 -0.0063 0.1490 0.9858 2.500 0.3041 0.00898 0.00214 -0.0070 0.0153 0.9894 2.750 0.3355 0.00948 0.00277 -0.0078 0.0086 0.9927 3.000 0.3670 0.00987 0.00324 -0.0087 0.0067 0.9958 3.250 0.3991 0.01049 0.00395 -0.0097 0.0057 0.9987 3.500 0.4243 0.01150 0.00506 -0.0093 0.0051 1.0000 3.750 0.4454 0.01226 0.00592 -0.0078 0.0044 1.0000 4.000 0.4672 0.01291 0.00667 -0.0065 0.0036 1.0000 4.250 0.4885 0.01378 0.00766 -0.0052 0.0031 1.0000 4.500 0.5089 0.01524 0.00930 -0.0036 0.0029 1.0000 4.750 0.5285 0.01764 0.01204 -0.0016 0.0028 1.0000 5.250 0.5670 0.02224 0.01733 0.0025 0.0029 1.0000 5.500 0.5841 0.02732 0.02299 0.0064 0.0036 1.0000 5.750 0.5982 0.03100 0.02704 0.0089 0.0041 1.0000 6.750 0.6434 0.04551 0.04258 0.0163 0.0055 1.0000 7.250 0.6553 0.05352 0.05095 0.0179 0.0053 1.0000 12.250 0.6523 0.14597 0.14359 -0.0234 0.0048 1.0000 12.500 0.6569 0.14972 0.14733 -0.0242 0.0049 1.0000