XFOIL Version 6.96 Calculated polar for: AH 85-L-120/17 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.500 -0.7020 0.02394 0.01822 -0.0433 0.9403 0.0147 -8.250 -0.6738 0.02111 0.01509 -0.0435 0.9382 0.0142 -8.000 -0.6556 0.02110 0.01506 -0.0425 0.9342 0.0156 -7.750 -0.6278 0.01951 0.01329 -0.0428 0.9320 0.0163 -7.500 -0.5995 0.01810 0.01175 -0.0432 0.9301 0.0164 -7.250 -0.5742 0.01698 0.01053 -0.0431 0.9280 0.0169 -7.000 -0.5544 0.01631 0.00978 -0.0419 0.9252 0.0179 -6.750 -0.5353 0.01573 0.00913 -0.0405 0.9220 0.0184 -6.500 -0.5161 0.01521 0.00855 -0.0391 0.9191 0.0191 -6.250 -0.5072 0.01398 0.00718 -0.0358 0.9159 0.0210 -6.000 -0.4904 0.01356 0.00675 -0.0340 0.9131 0.0231 -5.750 -0.4732 0.01317 0.00630 -0.0322 0.9101 0.0252 -5.500 -0.4546 0.01280 0.00588 -0.0305 0.9073 0.0281 -5.250 -0.4381 0.01227 0.00532 -0.0285 0.9045 0.0382 -5.000 -0.4243 0.01160 0.00490 -0.0261 0.9018 0.0872 -4.750 -0.4153 0.01080 0.00456 -0.0230 0.8987 0.1888 -4.500 -0.4083 0.00993 0.00424 -0.0195 0.8950 0.3110 -4.250 -0.4060 0.00896 0.00393 -0.0150 0.8913 0.4609 -4.000 -0.3995 0.00829 0.00374 -0.0108 0.8883 0.5797 -3.750 -0.3858 0.00796 0.00375 -0.0079 0.8856 0.6615 -3.500 -0.3654 0.00790 0.00383 -0.0063 0.8827 0.7083 -3.250 -0.3431 0.00800 0.00397 -0.0049 0.8799 0.7462 -3.000 -0.3193 0.00822 0.00420 -0.0038 0.8774 0.7716 -2.750 -0.2927 0.00838 0.00433 -0.0034 0.8753 0.7826 -2.500 -0.2663 0.00857 0.00448 -0.0029 0.8733 0.7921 -2.250 -0.2421 0.00871 0.00457 -0.0022 0.8703 0.8005 -2.000 -0.2148 0.00885 0.00471 -0.0020 0.8673 0.8064 -1.750 -0.1889 0.00901 0.00484 -0.0014 0.8644 0.8143 -1.500 -0.1607 0.00915 0.00497 -0.0013 0.8621 0.8199 -1.250 -0.1325 0.00933 0.00512 -0.0012 0.8602 0.8260 -1.000 -0.1077 0.00943 0.00520 -0.0006 0.8575 0.8318 -0.750 -0.0805 0.00949 0.00528 -0.0005 0.8541 0.8347 -0.500 -0.0533 0.00953 0.00532 -0.0004 0.8509 0.8382 -0.250 -0.0269 0.00953 0.00529 -0.0002 0.8482 0.8424 0.000 0.0000 0.00950 0.00523 0.0000 0.8458 0.8458 0.250 0.0269 0.00953 0.00528 0.0002 0.8424 0.8481 0.500 0.0533 0.00953 0.00532 0.0004 0.8381 0.8509 0.750 0.0806 0.00949 0.00528 0.0005 0.8347 0.8541 1.000 0.1078 0.00942 0.00518 0.0006 0.8317 0.8575 1.250 0.1326 0.00933 0.00513 0.0012 0.8261 0.8602 1.500 0.1608 0.00915 0.00497 0.0013 0.8198 0.8622 1.750 0.1890 0.00902 0.00485 0.0014 0.8143 0.8645 2.000 0.2148 0.00887 0.00473 0.0019 0.8068 0.8673 2.250 0.2421 0.00871 0.00457 0.0021 0.8006 0.8703 2.500 0.2664 0.00857 0.00448 0.0029 0.7922 0.8734 2.750 0.2928 0.00838 0.00433 0.0034 0.7826 0.8753 3.000 0.3197 0.00822 0.00421 0.0037 0.7726 0.8774 3.250 0.3432 0.00801 0.00401 0.0049 0.7486 0.8798 3.500 0.3658 0.00790 0.00386 0.0062 0.7123 0.8826 3.750 0.3852 0.00797 0.00374 0.0081 0.6568 0.8856 4.000 0.3978 0.00834 0.00375 0.0112 0.5696 0.8884 4.250 0.4056 0.00898 0.00393 0.0151 0.4579 0.8914 4.500 0.4102 0.00985 0.00423 0.0192 0.3233 0.8949 4.750 0.4170 0.01070 0.00455 0.0227 0.2023 0.8986 5.000 0.4262 0.01150 0.00487 0.0258 0.1006 0.9017 5.250 0.4378 0.01229 0.00534 0.0286 0.0365 0.9046 5.500 0.4554 0.01275 0.00582 0.0304 0.0288 0.9073 5.750 0.4731 0.01317 0.00631 0.0322 0.0256 0.9101 6.000 0.4916 0.01348 0.00664 0.0338 0.0223 0.9131 6.250 0.5043 0.01417 0.00738 0.0364 0.0199 0.9161 6.500 0.5173 0.01507 0.00840 0.0389 0.0188 0.9190 6.750 0.5363 0.01553 0.00893 0.0403 0.0183 0.9219 7.000 0.5547 0.01609 0.00956 0.0419 0.0170 0.9253 7.250 0.5744 0.01706 0.01061 0.0431 0.0171 0.9280 7.500 0.5993 0.01809 0.01174 0.0433 0.0163 0.9302 7.750 0.6277 0.01951 0.01329 0.0428 0.0162 0.9320 8.000 0.6574 0.02159 0.01560 0.0423 0.0158 0.9339 8.250 0.6752 0.02153 0.01554 0.0434 0.0142 0.9379 8.500 0.6988 0.02309 0.01727 0.0436 0.0140 0.9408 8.750 0.7217 0.02629 0.02087 0.0443 0.0145 0.9434