XFOIL Version 6.94 Calculated polar for: agora06 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0762 0.03619 0.02488 -0.0202 0.9988 0.3245 -2.750 -0.0546 0.03665 0.02540 -0.0197 0.9988 0.3462 -2.500 -0.0329 0.03716 0.02592 -0.0191 0.9988 0.3702 -2.250 -0.0114 0.03744 0.02644 -0.0186 0.9988 0.3981 -2.000 0.0099 0.03772 0.02699 -0.0180 0.9988 0.4326 -1.750 0.0302 0.03783 0.02756 -0.0169 0.9988 0.4735 -1.500 0.2320 0.03437 0.02402 -0.0435 0.8877 0.4937 -1.250 0.2460 0.03556 0.02533 -0.0376 0.8485 0.4998 -1.000 0.2507 0.03720 0.02707 -0.0292 0.8020 0.5072 -0.750 0.2496 0.03939 0.02931 -0.0191 0.7455 0.5160 -0.500 0.2472 0.04191 0.03191 -0.0088 0.6745 0.5266 -0.250 0.2516 0.04369 0.03383 -0.0006 0.5983 0.5440 0.000 0.3008 0.03839 0.02966 -0.0047 0.4498 0.6151 0.250 0.3146 0.03551 0.02695 0.0034 0.4219 1.0012 0.500 0.3381 0.03554 0.02603 0.0069 0.3960 1.0012 0.750 0.3638 0.03617 0.02592 0.0089 0.3734 1.0012 1.000 0.3907 0.03704 0.02627 0.0102 0.3548 1.0012 1.250 0.4172 0.03801 0.02682 0.0116 0.3425 1.0012 1.500 0.4446 0.03917 0.02769 0.0125 0.3306 1.0012 1.750 0.4721 0.04050 0.02877 0.0132 0.3189 1.0012 2.000 0.5010 0.04196 0.03008 0.0135 0.3095 1.0012 2.250 0.5304 0.04357 0.03162 0.0137 0.3033 1.0012 2.500 0.5601 0.04538 0.03336 0.0135 0.2966 1.0012 2.750 0.5900 0.04738 0.03535 0.0130 0.2892 1.0012 3.000 0.6215 0.04970 0.03780 0.0116 0.2835 1.0012 3.250 0.6533 0.05238 0.04063 0.0099 0.2815 1.0012 3.500 0.6873 0.05596 0.04459 0.0061 0.2805 1.0012 3.750 0.7200 0.06066 0.04972 0.0004 0.2790 1.0012 4.000 0.7488 0.06713 0.05664 -0.0076 0.2799 1.0012