XFOIL Version 6.96 Calculated polar for: GOE 491 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.750 -0.4900 0.10580 0.10216 -0.0033 1.0000 0.0325 -8.500 -0.4925 0.10391 0.10034 -0.0068 1.0000 0.0328 -8.250 -0.4928 0.10123 0.09770 -0.0090 1.0000 0.0329 -8.000 -0.4901 0.09824 0.09474 -0.0117 1.0000 0.0330 -7.750 -0.4836 0.09486 0.09136 -0.0145 1.0000 0.0330 -7.500 -0.4810 0.08945 0.08598 -0.0158 1.0000 0.0334 -7.250 -0.4773 0.08438 0.08095 -0.0122 1.0000 0.0342 -7.000 -0.4703 0.08096 0.07755 -0.0117 1.0000 0.0349 -6.750 -0.4630 0.07765 0.07424 -0.0122 1.0000 0.0358 -6.500 -0.4545 0.07438 0.07095 -0.0134 1.0000 0.0368 -6.250 -0.4447 0.07101 0.06756 -0.0149 1.0000 0.0381 -6.000 -0.4331 0.06761 0.06413 -0.0166 1.0000 0.0395 -5.750 -0.4180 0.06418 0.06064 -0.0188 1.0000 0.0417 -5.500 -0.3879 0.06183 0.05800 -0.0237 1.0000 0.0439 -5.250 -0.3700 0.05872 0.05472 -0.0242 1.0000 0.0442 -5.000 -0.3613 0.05281 0.04874 -0.0243 1.0000 0.0448 -4.750 -0.3521 0.04890 0.04487 -0.0232 1.0000 0.0458 -4.500 -0.3391 0.04602 0.04198 -0.0221 1.0000 0.0470 -4.250 -0.3237 0.04344 0.03934 -0.0212 1.0000 0.0496 -4.000 -0.2925 0.04258 0.03796 -0.0209 1.0000 0.0569 -3.750 -0.2807 0.03745 0.03266 -0.0198 1.0000 0.0581 -3.500 -0.2668 0.03438 0.02958 -0.0185 1.0000 0.0594 -3.250 -0.2505 0.03203 0.02714 -0.0169 1.0000 0.0610 -3.000 -0.2282 0.02620 0.02061 -0.0141 1.0000 0.0424 -2.750 -0.2099 0.02287 0.01690 -0.0120 1.0000 0.0406 -2.500 -0.1889 0.02032 0.01387 -0.0099 1.0000 0.0424 -2.250 -0.1670 0.01804 0.01107 -0.0081 1.0000 0.0450 -2.000 -0.1457 0.01664 0.00954 -0.0068 1.0000 0.0486 -1.750 -0.1232 0.01580 0.00844 -0.0053 1.0000 0.0555 -1.500 -0.1017 0.01515 0.00771 -0.0042 1.0000 0.0665 -1.250 -0.0796 0.01444 0.00687 -0.0030 1.0000 0.0771 -1.000 -0.0460 0.01409 0.00647 -0.0045 0.9964 0.0909 -0.750 -0.0020 0.01372 0.00595 -0.0078 0.9891 0.0986 -0.500 0.0411 0.01282 0.00514 -0.0110 0.9824 0.1055 -0.250 0.0830 0.01228 0.00460 -0.0138 0.9729 0.1110 0.000 0.1230 0.01165 0.00407 -0.0162 0.9623 0.1173 0.250 0.1638 0.01112 0.00364 -0.0188 0.9509 0.1297 0.500 0.2050 0.01057 0.00324 -0.0213 0.9389 0.1519 0.750 0.3066 0.00827 0.00319 -0.0371 0.9437 1.0000 1.000 0.3601 0.00806 0.00290 -0.0421 0.9196 1.0000 1.250 0.4183 0.00786 0.00256 -0.0480 0.8736 1.0000 1.500 0.4502 0.00794 0.00242 -0.0481 0.8040 1.0000 1.750 0.4733 0.00818 0.00235 -0.0464 0.7288 1.0000 2.000 0.4940 0.00854 0.00236 -0.0444 0.6619 1.0000 2.250 0.5149 0.00891 0.00244 -0.0427 0.6084 1.0000 2.500 0.5364 0.00925 0.00256 -0.0412 0.5658 1.0000 2.750 0.5576 0.00960 0.00267 -0.0397 0.5236 1.0000 3.000 0.5792 0.00993 0.00284 -0.0383 0.4868 1.0000 3.250 0.6018 0.01023 0.00302 -0.0372 0.4614 1.0000 3.500 0.6247 0.01052 0.00323 -0.0361 0.4397 1.0000 3.750 0.6478 0.01079 0.00346 -0.0351 0.4177 1.0000 4.000 0.6707 0.01106 0.00371 -0.0340 0.3953 1.0000 4.250 0.6938 0.01127 0.00393 -0.0330 0.3693 1.0000 4.500 0.7167 0.01150 0.00415 -0.0319 0.3394 1.0000 4.750 0.7392 0.01177 0.00436 -0.0308 0.2963 1.0000 5.000 0.7591 0.01240 0.00465 -0.0294 0.2150 1.0000 5.250 0.7744 0.01385 0.00541 -0.0275 0.0932 1.0000 5.500 0.7943 0.01476 0.00621 -0.0259 0.0714 1.0000 5.750 0.8148 0.01553 0.00702 -0.0245 0.0629 1.0000 6.000 0.8342 0.01642 0.00794 -0.0229 0.0563 1.0000 6.250 0.8531 0.01744 0.00905 -0.0212 0.0523 1.0000 6.500 0.8724 0.01841 0.01005 -0.0196 0.0467 1.0000 6.750 0.8878 0.02055 0.01222 -0.0175 0.0431 1.0000 7.000 0.9090 0.02173 0.01358 -0.0160 0.0405 1.0000 7.250 0.9291 0.02308 0.01505 -0.0145 0.0364 1.0000 7.500 0.9488 0.02506 0.01715 -0.0131 0.0343 1.0000 7.750 0.9675 0.02805 0.02034 -0.0116 0.0328 1.0000 8.000 0.9828 0.03207 0.02476 -0.0097 0.0320 1.0000 8.250 0.9958 0.03396 0.02712 -0.0074 0.0302 1.0000 8.500 1.0086 0.03633 0.02994 -0.0051 0.0292 1.0000 8.750 1.0160 0.04004 0.03408 -0.0026 0.0295 1.0000 9.000 1.0177 0.04463 0.03909 0.0000 0.0302 1.0000