XFOIL Version 6.94 Calculated polar for: RAF 32 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 0.5993 0.01010 0.00429 -0.1431 0.7573 0.6481 0.500 0.6480 0.00997 0.00442 -0.1414 0.7460 0.7385 1.000 0.7339 0.00949 0.00438 -0.1476 0.7351 0.9999 1.500 0.7866 0.00971 0.00446 -0.1471 0.7240 1.0000 2.000 0.8381 0.00987 0.00452 -0.1463 0.7120 1.0000 2.500 0.8785 0.00969 0.00410 -0.1424 0.6746 1.0000 3.000 0.9093 0.00985 0.00396 -0.1366 0.6155 1.0000 3.500 0.9353 0.01033 0.00412 -0.1303 0.5489 1.0000 4.000 0.9405 0.01138 0.00456 -0.1201 0.4429 1.0000 4.500 0.9453 0.01284 0.00545 -0.1106 0.3405 1.0000 5.000 0.9269 0.01583 0.00734 -0.0983 0.1490 1.0000 5.500 0.9350 0.01802 0.00902 -0.0910 0.0472 1.0000 6.000 0.9610 0.01943 0.01030 -0.0866 0.0092 1.0000 6.500 0.9944 0.02051 0.01148 -0.0833 0.0095 1.0000 7.000 1.0273 0.02167 0.01276 -0.0801 0.0105 1.0000 7.500 1.0571 0.02307 0.01431 -0.0766 0.0115 1.0000 8.000 1.0867 0.02456 0.01594 -0.0733 0.0130 1.0000 8.500 1.1117 0.02645 0.01801 -0.0696 0.0139 1.0000 9.000 1.1343 0.02862 0.02036 -0.0658 0.0146 1.0000 9.500 1.1492 0.03148 0.02341 -0.0617 0.0151 1.0000 10.000 1.1571 0.03506 0.02713 -0.0575 0.0157 1.0000 10.500 1.1601 0.03931 0.03146 -0.0534 0.0160 1.0000 11.000 1.1788 0.04248 0.03480 -0.0504 0.0171 1.0000 11.500 1.2032 0.04562 0.03792 -0.0470 0.0187 1.0000 12.000 1.2665 0.04698 0.03949 -0.0448 0.0223 1.0000 12.500 1.4124 0.05273 0.04579 -0.0493 0.0253 1.0000 13.000 1.4244 0.05964 0.05327 -0.0455 0.0253 1.0000 13.500 1.4026 0.06635 0.06054 -0.0400 0.0253 1.0000 14.000 1.3808 0.07410 0.06876 -0.0362 0.0253 1.0000 14.500 1.3437 0.08250 0.07762 -0.0334 0.0253 1.0000 15.000 1.3143 0.09200 0.08749 -0.0328 0.0252 1.0000