XFOIL Version 6.94 Calculated polar for: GOE 235 (SCHTTE-LANZ) 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 ------- -------- --------- --------- -------- ------- ------- 1.000 0.3804 0.01150 0.00423 -0.0503 0.5978 0.0564 1.500 0.4362 0.01104 0.00351 -0.0496 0.5200 0.0550 2.000 0.4893 0.01151 0.00331 -0.0491 0.3634 0.0544 2.500 0.5423 0.01213 0.00338 -0.0488 0.2334 0.0609 3.000 0.5975 0.01219 0.00334 -0.0486 0.2111 0.0678 3.500 0.6532 0.01231 0.00350 -0.0483 0.1997 0.0761 4.000 0.7044 0.01055 0.00382 -0.0475 0.1845 1.0000 4.500 0.7603 0.01087 0.00403 -0.0471 0.1729 1.0000 5.000 0.8155 0.01123 0.00428 -0.0467 0.1529 1.0000 5.500 0.8645 0.01293 0.00532 -0.0458 0.0431 1.0000 6.000 0.9173 0.01372 0.00613 -0.0450 0.0398 1.0000 6.500 0.9684 0.01472 0.00713 -0.0442 0.0382 1.0000 7.000 1.0186 0.01584 0.00828 -0.0432 0.0358 1.0000 7.500 1.0656 0.01734 0.00981 -0.0418 0.0338 1.0000 8.000 1.1114 0.01888 0.01141 -0.0403 0.0327 1.0000 8.500 1.1567 0.02037 0.01296 -0.0387 0.0321 1.0000 9.000 1.1974 0.02249 0.01506 -0.0368 0.0303 1.0000 9.500 1.2399 0.02450 0.01711 -0.0350 0.0303 1.0000 10.000 1.2810 0.02703 0.01975 -0.0331 0.0295 1.0000 10.500 1.3212 0.02916 0.02210 -0.0311 0.0288 1.0000 11.000 1.3589 0.03189 0.02505 -0.0290 0.0281 1.0000 11.500 1.3933 0.03490 0.02828 -0.0267 0.0275 1.0000 12.000 1.4251 0.03813 0.03169 -0.0244 0.0269 1.0000 12.500 1.4532 0.04213 0.03589 -0.0220 0.0266 1.0000 13.000 1.4757 0.04746 0.04144 -0.0198 0.0261 1.0000 13.500 1.4656 0.05349 0.04792 -0.0148 0.0259 1.0000 14.000 1.4442 0.05924 0.05406 -0.0106 0.0258 1.0000 14.500 1.4281 0.06600 0.06114 -0.0090 0.0260 1.0000 15.000 1.3932 0.07399 0.06952 -0.0103 0.0261 1.0000