XFOIL Version 6.94 Calculated polar for: RG 12A-1.8/9.0 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 ------- -------- --------- --------- -------- ------- ------- 0.000 0.2116 0.00756 0.00317 -0.0433 0.8850 0.9240 0.500 0.2718 0.00746 0.00299 -0.0430 0.8509 0.9460 1.000 0.3428 0.00749 0.00287 -0.0454 0.8107 0.9608 1.500 0.4103 0.00759 0.00280 -0.0476 0.7638 0.9714 2.000 0.4836 0.00778 0.00277 -0.0513 0.7123 0.9781 2.500 0.5489 0.00802 0.00283 -0.0535 0.6581 0.9882 3.000 0.6188 0.00831 0.00298 -0.0569 0.6000 0.9974 3.500 0.6554 0.00861 0.00314 -0.0535 0.5461 1.0000 4.000 0.6871 0.00903 0.00338 -0.0491 0.4927 1.0000 4.500 0.7310 0.00961 0.00376 -0.0470 0.4303 1.0000 5.000 0.7768 0.01031 0.00426 -0.0454 0.3579 1.0000 5.500 0.8206 0.01129 0.00487 -0.0436 0.2674 1.0000 6.000 0.8612 0.01273 0.00571 -0.0416 0.1546 1.0000 6.500 0.9024 0.01423 0.00676 -0.0397 0.0750 1.0000 7.000 0.9419 0.01594 0.00819 -0.0374 0.0280 1.0000 7.500 0.9844 0.01726 0.00958 -0.0354 0.0205 1.0000 8.000 1.0261 0.01853 0.01094 -0.0335 0.0159 1.0000 8.500 1.0663 0.01984 0.01240 -0.0314 0.0130 1.0000 9.000 1.1038 0.02129 0.01407 -0.0289 0.0115 1.0000 9.500 1.1345 0.02316 0.01616 -0.0255 0.0106 1.0000 10.000 1.1609 0.02514 0.01847 -0.0214 0.0096 1.0000 10.500 1.1864 0.02670 0.02030 -0.0175 0.0073 1.0000 11.000 1.2046 0.02876 0.02267 -0.0128 0.0055 1.0000 12.500 1.1946 0.04125 0.03633 0.0006 0.0029 1.0000 13.000 1.1758 0.04833 0.04382 0.0012 0.0027 1.0000 13.500 1.1514 0.05758 0.05347 -0.0014 0.0027 1.0000 14.000 1.1186 0.07002 0.06632 -0.0075 0.0027 1.0000 14.500 1.0808 0.08555 0.08222 -0.0166 0.0027 1.0000 15.000 1.0363 0.10474 0.10174 -0.0281 0.0030 1.0000 15.500 0.9952 0.12525 0.12247 -0.0401 0.0033 1.0000 16.000 0.9609 0.14541 0.14275 -0.0514 0.0036 1.0000