XFOIL Version 6.94 Calculated polar for: AIRFOIL PROFILE12A 9.00% 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.2552 0.00610 0.00205 -0.0536 0.8436 0.8732 0.500 0.3056 0.00620 0.00202 -0.0517 0.7995 0.8922 1.000 0.3524 0.00635 0.00202 -0.0491 0.7528 0.9128 1.500 0.3990 0.00652 0.00204 -0.0466 0.7062 0.9305 2.000 0.4473 0.00671 0.00206 -0.0447 0.6596 0.9456 2.500 0.5000 0.00692 0.00214 -0.0439 0.6116 0.9612 3.000 0.5620 0.00722 0.00229 -0.0454 0.5570 0.9757 3.500 0.6303 0.00763 0.00251 -0.0484 0.4967 0.9897 4.000 0.6866 0.00811 0.00280 -0.0492 0.4308 1.0000 4.500 0.7315 0.00883 0.00316 -0.0476 0.3463 1.0000 5.000 0.7800 0.00958 0.00366 -0.0466 0.2754 1.0000 5.500 0.8274 0.01053 0.00428 -0.0456 0.1967 1.0000 6.000 0.8737 0.01164 0.00506 -0.0443 0.1230 1.0000 6.500 0.9156 0.01326 0.00625 -0.0424 0.0430 1.0000 7.000 0.9596 0.01470 0.00772 -0.0404 0.0257 1.0000 7.500 0.9966 0.01684 0.01009 -0.0372 0.0214 1.0000 8.000 1.0382 0.01836 0.01183 -0.0350 0.0198 1.0000 8.500 1.0754 0.02046 0.01416 -0.0321 0.0182 1.0000 9.000 1.1120 0.02251 0.01639 -0.0295 0.0162 1.0000 9.500 1.1325 0.02851 0.02284 -0.0252 0.0138 1.0000 10.000 1.1632 0.03066 0.02536 -0.0219 0.0129 1.0000 10.500 1.1820 0.03374 0.02885 -0.0174 0.0113 1.0000 11.000 1.1925 0.03551 0.03076 -0.0122 0.0098 1.0000 11.500 1.1832 0.04030 0.03586 -0.0068 0.0090 1.0000 12.000 1.1562 0.04759 0.04364 -0.0030 0.0088 1.0000 12.500 1.1161 0.05754 0.05406 -0.0034 0.0087 1.0000 13.000 1.0821 0.06876 0.06566 -0.0083 0.0086 1.0000 13.500 1.0526 0.08163 0.07884 -0.0162 0.0087 1.0000 14.000 1.0205 0.09739 0.09488 -0.0268 0.0087 1.0000 14.500 0.7348 0.13330 0.13158 -0.0519 0.0149 1.0000