XFOIL Version 6.94 Calculated polar for: GOE 693 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 1.500 0.5868 0.00907 0.00352 -0.0925 0.6953 1.0000 2.000 0.6378 0.00918 0.00362 -0.0915 0.6803 1.0000 2.500 0.6896 0.00935 0.00371 -0.0905 0.6642 1.0000 3.000 0.7399 0.00944 0.00371 -0.0891 0.6373 1.0000 3.500 0.7860 0.00960 0.00357 -0.0868 0.5794 1.0000 4.000 0.8252 0.01034 0.00370 -0.0834 0.4639 1.0000 4.500 0.8668 0.01127 0.00421 -0.0809 0.3880 1.0000 5.000 0.9040 0.01256 0.00496 -0.0779 0.2991 1.0000 5.500 0.9464 0.01351 0.00561 -0.0758 0.2336 1.0000 6.000 0.9790 0.01517 0.00670 -0.0723 0.1475 1.0000 7.000 1.0374 0.01867 0.00939 -0.0641 0.0065 1.0000 7.500 1.0733 0.01964 0.01046 -0.0609 0.0055 1.0000 8.000 1.1067 0.02074 0.01168 -0.0574 0.0053 1.0000 8.500 1.1382 0.02205 0.01315 -0.0538 0.0053 1.0000 9.000 1.1672 0.02360 0.01486 -0.0503 0.0053 1.0000 9.500 1.1936 0.02542 0.01687 -0.0467 0.0053 1.0000 10.000 1.2174 0.02753 0.01917 -0.0433 0.0054 1.0000 10.500 1.2382 0.03001 0.02186 -0.0400 0.0056 1.0000 11.000 1.2544 0.03303 0.02508 -0.0368 0.0057 1.0000 11.500 1.2648 0.03677 0.02905 -0.0340 0.0059 1.0000 12.000 1.2691 0.04133 0.03382 -0.0314 0.0061 1.0000 12.500 1.2690 0.04659 0.03929 -0.0294 0.0063 1.0000 13.000 1.2655 0.05250 0.04540 -0.0279 0.0065 1.0000 13.500 1.2608 0.05880 0.05188 -0.0268 0.0067 1.0000 14.000 1.2571 0.06509 0.05828 -0.0257 0.0069 1.0000 14.500 1.2620 0.07067 0.06403 -0.0252 0.0071 1.0000 15.000 1.2675 0.07665 0.07033 -0.0253 0.0075 1.0000 15.500 1.2697 0.08278 0.07675 -0.0245 0.0081 1.0000 16.000 1.2737 0.08859 0.08277 -0.0233 0.0087 1.0000 16.500 1.2775 0.09493 0.08934 -0.0236 0.0091 1.0000 17.000 1.2480 0.10787 0.10300 -0.0282 0.0101 1.0000 17.500 1.2290 0.11902 0.11450 -0.0323 0.0107 1.0000 19.000 1.1374 0.16747 0.16414 -0.0622 0.0117 1.0000 19.500 1.1052 0.18772 0.18466 -0.0760 0.0116 1.0000