XFOIL Version 6.94 Calculated polar for: DFVLR R-4 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 0.4293 0.00906 0.00458 -0.1115 0.8646 0.7105 0.500 0.4843 0.00908 0.00461 -0.1107 0.8454 0.7159 1.000 0.5410 0.00912 0.00465 -0.1104 0.8221 0.7234 1.500 0.5980 0.00910 0.00458 -0.1103 0.7867 0.7296 2.000 0.6483 0.00936 0.00465 -0.1084 0.7074 0.7341 3.000 0.7169 0.01325 0.00612 -0.1006 0.2149 0.7461 3.500 0.7638 0.01426 0.00674 -0.0992 0.1433 0.7515 4.000 0.8115 0.01501 0.00742 -0.0975 0.1209 0.7564 4.500 0.8616 0.01572 0.00809 -0.0965 0.1089 0.7613 5.000 0.9117 0.01655 0.00887 -0.0957 0.0999 0.7671 5.500 0.9620 0.01715 0.00954 -0.0948 0.0942 0.7717 6.000 1.0075 0.01806 0.01047 -0.0930 0.0890 0.7753 6.500 1.0557 0.01877 0.01128 -0.0917 0.0849 0.7793 7.000 1.1008 0.01997 0.01246 -0.0901 0.0807 0.7839 7.500 1.1503 0.02068 0.01330 -0.0892 0.0768 0.7886 8.000 1.1952 0.02169 0.01431 -0.0876 0.0720 0.7921 8.500 1.2393 0.02228 0.01509 -0.0856 0.0675 0.7960 9.000 1.2803 0.02319 0.01608 -0.0833 0.0628 0.8001 9.500 1.3159 0.02410 0.01701 -0.0801 0.0583 0.8048 10.500 1.3917 0.02603 0.01924 -0.0750 0.0495 0.8121 11.000 1.4254 0.02721 0.02050 -0.0720 0.0460 0.8156 12.000 1.4932 0.02975 0.02343 -0.0665 0.0384 0.8237 12.500 1.5224 0.03148 0.02527 -0.0636 0.0347 0.8273 13.000 1.5457 0.03374 0.02764 -0.0603 0.0310 0.8304 13.500 1.5665 0.03618 0.03032 -0.0570 0.0277 0.8339 14.000 1.5776 0.03950 0.03388 -0.0532 0.0251 0.8375 14.500 1.5787 0.04385 0.03841 -0.0494 0.0230 0.8412 15.000 1.5840 0.04821 0.04310 -0.0467 0.0210 0.8455 15.500 1.5756 0.05424 0.04934 -0.0446 0.0198 0.8490 16.000 1.5657 0.06120 0.05665 -0.0439 0.0188 0.8521 16.500 1.5464 0.07022 0.06602 -0.0453 0.0181 0.8556 17.000 1.5175 0.08222 0.07836 -0.0497 0.0176 0.8591 17.500 1.4735 0.09845 0.09497 -0.0580 0.0174 0.8620 18.000 1.4152 0.11867 0.11558 -0.0695 0.0175 0.8640 18.500 1.3550 0.14001 0.13725 -0.0823 0.0177 0.8655 19.000 1.3058 0.16000 0.15749 -0.0953 0.0175 0.8672 19.500 1.2699 0.17827 0.17596 -0.1081 0.0173 0.8695 20.000 1.2427 0.19524 0.19309 -0.1203 0.0169 0.8722