XFOIL Version 6.96 Calculated polar for: B-29 ROOT 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 ------ -------- --------- --------- -------- -------- -------- -17.500 -0.8667 0.07932 0.07404 -0.0761 1.0000 0.0821 -17.250 -0.9110 0.07154 0.06599 -0.0792 1.0000 0.0822 -17.000 -0.9366 0.06642 0.06065 -0.0806 1.0000 0.0827 -16.750 -0.9565 0.06222 0.05627 -0.0813 1.0000 0.0833 -16.500 -0.9735 0.05856 0.05241 -0.0814 1.0000 0.0839 -16.250 -0.9886 0.05526 0.04892 -0.0810 1.0000 0.0845 -16.000 -1.0024 0.05234 0.04580 -0.0802 1.0000 0.0852 -15.750 -1.0172 0.04970 0.04294 -0.0789 1.0000 0.0860 -15.500 -1.0294 0.04741 0.04044 -0.0772 1.0000 0.0865 -15.250 -1.0203 0.04555 0.03851 -0.0765 1.0000 0.0876 -15.000 -1.0001 0.04444 0.03747 -0.0762 1.0000 0.0889 -14.750 -0.9903 0.04319 0.03620 -0.0753 1.0000 0.0900 -14.500 -0.9839 0.04185 0.03480 -0.0741 1.0000 0.0911 -14.250 -0.9781 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4.750 0.7307 0.02092 0.01410 -0.0277 0.5157 0.9152 5.000 0.7751 0.02113 0.01431 -0.0310 0.5114 0.9183 5.250 0.8189 0.02125 0.01437 -0.0341 0.5079 0.9227 5.500 0.8440 0.02152 0.01459 -0.0337 0.5048 0.9320 5.750 0.8807 0.02199 0.01522 -0.0361 0.5000 0.9360 6.000 0.9139 0.02229 0.01559 -0.0375 0.4956 0.9424 6.250 0.9439 0.02242 0.01571 -0.0382 0.4918 0.9498 6.500 0.9914 0.02244 0.01566 -0.0421 0.4883 0.9529 6.750 1.0259 0.02278 0.01605 -0.0439 0.4846 0.9589 7.000 1.0508 0.02321 0.01663 -0.0442 0.4799 0.9664 7.250 1.0887 0.02334 0.01682 -0.0467 0.4753 0.9717 7.500 1.1228 0.02334 0.01677 -0.0481 0.4713 0.9785 7.750 1.1683 0.02322 0.01661 -0.0518 0.4664 0.9818 8.000 1.1968 0.02336 0.01692 -0.0529 0.4596 0.9889 8.250 1.2380 0.02295 0.01645 -0.0556 0.4532 0.9936 8.500 1.2758 0.02266 0.01616 -0.0581 0.4454 0.9988 8.750 1.2885 0.02262 0.01617 -0.0559 0.4383 1.0000 9.000 1.3070 0.02246 0.01592 -0.0544 0.4330 1.0000 9.250 1.3133 0.02273 0.01626 -0.0511 0.4284 1.0000 9.500 1.3111 0.02308 0.01671 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14.500 1.2026 0.03985 0.03395 0.0332 0.2938 1.0000 14.750 1.1875 0.04276 0.03692 0.0351 0.2812 1.0000 15.000 1.1762 0.04550 0.03964 0.0365 0.2678 1.0000 15.250 1.1651 0.04835 0.04245 0.0378 0.2533 1.0000 15.500 1.1521 0.05154 0.04558 0.0389 0.2376 1.0000 15.750 1.1382 0.05493 0.04888 0.0399 0.2216 1.0000 16.000 1.1251 0.05837 0.05223 0.0406 0.2065 1.0000 16.250 1.1129 0.06183 0.05559 0.0412 0.1927 1.0000 16.500 1.1025 0.06517 0.05883 0.0417 0.1804 1.0000 16.750 1.0955 0.06818 0.06171 0.0421 0.1699 1.0000 17.000 1.0898 0.07118 0.06464 0.0424 0.1605 1.0000 17.250 1.0866 0.07398 0.06742 0.0426 0.1519 1.0000 17.500 1.0870 0.07624 0.06953 0.0430 0.1445 1.0000 17.750 1.0843 0.07920 0.07256 0.0428 0.1379 1.0000 18.000 1.0885 0.08108 0.07430 0.0431 0.1320 1.0000 18.250 1.0877 0.08390 0.07721 0.0429 0.1267 1.0000 18.500 1.0924 0.08582 0.07903 0.0430 0.1219 1.0000 18.750 1.0945 0.08826 0.08152 0.0428 0.1176 1.0000 19.000 1.0972 0.09062 0.08391 0.0426 0.1137 1.0000 19.250 1.1076 0.09178 0.08493 0.0428 0.1098 1.0000