XFOIL Version 6.94 Calculated polar for: GOE 344 (PFALZ 71) 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 0.2744 0.00801 0.00210 -0.0521 0.9137 0.1978 0.500 0.3337 0.00755 0.00196 -0.0524 0.8500 0.3356 1.000 0.4211 0.00665 0.00221 -0.0583 0.7085 0.9988 1.500 0.4729 0.00716 0.00225 -0.0572 0.6187 1.0000 2.000 0.5206 0.00765 0.00234 -0.0552 0.5340 1.0000 2.500 0.5679 0.00818 0.00250 -0.0534 0.4485 1.0000 3.000 0.6119 0.00913 0.00272 -0.0511 0.3184 1.0000 3.500 0.6584 0.00985 0.00294 -0.0494 0.2305 1.0000 4.000 0.7049 0.01061 0.00329 -0.0476 0.1495 1.0000 4.500 0.7507 0.01154 0.00381 -0.0457 0.0784 1.0000 5.000 0.7944 0.01291 0.00466 -0.0435 0.0044 1.0000 5.500 0.8445 0.01338 0.00517 -0.0422 0.0039 1.0000 6.000 0.8944 0.01400 0.00587 -0.0409 0.0038 1.0000 6.500 0.9438 0.01472 0.00679 -0.0394 0.0039 1.0000 7.000 0.9928 0.01553 0.00780 -0.0380 0.0041 1.0000 7.500 1.0405 0.01653 0.00902 -0.0363 0.0043 1.0000 8.000 1.0867 0.01774 0.01047 -0.0343 0.0048 1.0000 8.500 1.1310 0.01914 0.01210 -0.0321 0.0053 1.0000 9.000 1.1704 0.02104 0.01436 -0.0290 0.0060 1.0000 9.500 1.2014 0.02378 0.01748 -0.0248 0.0067 1.0000 10.000 1.2182 0.02811 0.02229 -0.0188 0.0073 1.0000 10.500 1.2460 0.03061 0.02506 -0.0147 0.0081 1.0000 11.000 1.2546 0.03519 0.03019 -0.0082 0.0097 1.0000 11.500 1.2290 0.04264 0.03831 0.0005 0.0110 1.0000 12.000 1.1932 0.05128 0.04751 0.0046 0.0118 1.0000 12.500 1.1564 0.06228 0.05896 0.0014 0.0120 1.0000 13.000 1.1217 0.07635 0.07341 -0.0079 0.0120 1.0000 13.500 1.0887 0.09243 0.08977 -0.0191 0.0116 1.0000 14.000 1.0535 0.10984 0.10739 -0.0302 0.0111 1.0000 14.500 1.0225 0.12699 0.12467 -0.0401 0.0103 1.0000 15.000 0.9911 0.14591 0.14366 -0.0503 0.0094 1.0000