XFOIL Version 6.96 Calculated polar for: EPPLER E1212 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -5.000 -0.2239 0.01177 0.00482 -0.0636 0.5840 0.2045 -4.500 -0.1679 0.01171 0.00464 -0.0630 0.5629 0.2113 -4.000 -0.1123 0.01153 0.00433 -0.0623 0.5428 0.2178 -3.500 -0.0566 0.01141 0.00415 -0.0617 0.5238 0.2247 -3.000 -0.0003 0.01140 0.00400 -0.0611 0.5055 0.2309 -2.500 0.0550 0.01121 0.00376 -0.0605 0.4879 0.2385 -2.000 0.1113 0.01115 0.00365 -0.0600 0.4714 0.2457 -1.500 0.1675 0.01108 0.00351 -0.0594 0.4554 0.2532 -1.000 0.2233 0.01102 0.00344 -0.0589 0.4405 0.2620 -0.500 0.2793 0.01107 0.00340 -0.0584 0.4257 0.2702 0.000 0.3344 0.01105 0.00336 -0.0577 0.4113 0.2818 0.500 0.3905 0.01104 0.00336 -0.0573 0.3979 0.2938 1.000 0.4460 0.01110 0.00342 -0.0567 0.3851 0.3083 2.000 0.5558 0.01116 0.00360 -0.0555 0.3610 0.3587 2.500 0.6100 0.01120 0.00378 -0.0549 0.3500 0.4070 3.000 0.6630 0.01121 0.00399 -0.0541 0.3394 0.4835 3.500 0.7155 0.01111 0.00426 -0.0532 0.3296 0.5956 4.000 0.7623 0.01102 0.00461 -0.0510 0.3193 0.7617 4.500 0.8427 0.01089 0.00494 -0.0552 0.3097 1.0000 5.000 0.8915 0.01133 0.00526 -0.0537 0.3008 1.0000 5.500 0.9426 0.01164 0.00554 -0.0525 0.2924 1.0000 6.000 0.9914 0.01215 0.00596 -0.0511 0.2842 1.0000 6.500 1.0424 0.01250 0.00632 -0.0500 0.2769 1.0000 7.000 1.0896 0.01311 0.00682 -0.0484 0.2685 1.0000 7.500 1.1398 0.01347 0.00724 -0.0473 0.2621 1.0000 8.000 1.1862 0.01405 0.00776 -0.0456 0.2551 1.0000 8.500 1.2325 0.01463 0.00836 -0.0440 0.2489 1.0000 9.000 1.2775 0.01517 0.00892 -0.0422 0.2424 1.0000 9.500 1.3179 0.01599 0.00969 -0.0399 0.2362 1.0000 10.000 1.3595 0.01655 0.01036 -0.0376 0.2314 1.0000