XFOIL Version 6.94 Calculated polar for: NASA SC(2)-0402 AIRFOIL 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 0.3161 0.00601 0.00193 -0.0646 0.7953 0.8011 0.500 0.3665 0.00633 0.00197 -0.0624 0.7020 0.8426 1.000 0.4158 0.00663 0.00202 -0.0602 0.6128 0.8856 1.500 0.4589 0.00685 0.00198 -0.0566 0.5089 1.0000 2.000 0.5156 0.00776 0.00223 -0.0570 0.3677 1.0000 2.500 0.5708 0.00886 0.00265 -0.0572 0.2218 1.0000 3.000 0.6251 0.01015 0.00325 -0.0573 0.0989 1.0000 3.500 0.6803 0.01111 0.00398 -0.0571 0.0665 1.0000 4.000 0.7356 0.01196 0.00476 -0.0568 0.0554 1.0000 4.500 0.7900 0.01288 0.00563 -0.0565 0.0489 1.0000 5.000 0.8435 0.01391 0.00666 -0.0559 0.0446 1.0000 5.500 0.8955 0.01518 0.00793 -0.0552 0.0420 1.0000 6.000 0.9475 0.01643 0.00923 -0.0544 0.0398 1.0000 6.500 0.9977 0.01809 0.01091 -0.0534 0.0379 1.0000 7.000 1.0485 0.01965 0.01263 -0.0524 0.0364 1.0000 7.500 1.0981 0.02143 0.01454 -0.0514 0.0352 1.0000 8.000 1.1457 0.02394 0.01713 -0.0503 0.0342 1.0000 8.500 1.1906 0.02693 0.02053 -0.0488 0.0336 1.0000 9.000 1.2316 0.03006 0.02417 -0.0470 0.0326 1.0000 9.500 1.2779 0.03034 0.02430 -0.0462 0.0298 1.0000 10.000 1.3119 0.03347 0.02803 -0.0439 0.0286 1.0000 10.500 1.3477 0.03559 0.03039 -0.0419 0.0274 1.0000 11.000 1.3817 0.03759 0.03249 -0.0400 0.0266 1.0000 11.500 1.3940 0.04157 0.03702 -0.0365 0.0252 1.0000 12.000 1.3881 0.04638 0.04228 -0.0317 0.0250 1.0000 12.500 1.3375 0.05657 0.05307 -0.0300 0.0253 1.0000 13.000 1.2586 0.07670 0.07378 -0.0429 0.0256 1.0000 13.500 1.4234 0.05525 0.05152 -0.0305 0.0217 1.0000 14.000 1.3491 0.07554 0.07241 -0.0429 0.0222 1.0000