XFOIL Version 6.94 Calculated polar for: NACA CYH 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 0.3403 0.01126 0.00303 -0.0356 0.6295 0.0454 0.500 0.3937 0.01136 0.00313 -0.0355 0.5984 0.0639 1.000 0.4449 0.00889 0.00347 -0.0348 0.5909 0.9277 1.500 0.5210 0.00905 0.00356 -0.0391 0.5671 0.9635 2.000 0.6065 0.00939 0.00348 -0.0458 0.5053 0.9889 2.500 0.6825 0.00978 0.00351 -0.0508 0.4483 1.0000 3.000 0.7269 0.01035 0.00371 -0.0496 0.3830 1.0000 3.500 0.7717 0.01090 0.00399 -0.0484 0.3570 1.0000 4.000 0.8167 0.01147 0.00445 -0.0473 0.3376 1.0000 4.500 0.8630 0.01204 0.00506 -0.0463 0.3232 1.0000 5.000 0.9095 0.01269 0.00582 -0.0454 0.3123 1.0000 5.500 0.9550 0.01338 0.00662 -0.0444 0.2982 1.0000 6.500 1.0460 0.01311 0.00619 -0.0422 0.2745 1.0000 7.000 1.0778 0.01416 0.00675 -0.0390 0.2070 1.0000 7.500 1.1042 0.01571 0.00766 -0.0353 0.1482 1.0000 8.000 1.1361 0.01712 0.00876 -0.0328 0.1110 1.0000 8.500 1.1655 0.01878 0.01004 -0.0301 0.0759 1.0000 9.000 1.1539 0.02330 0.01492 -0.0210 0.0173 1.0000 9.500 1.1600 0.02647 0.01853 -0.0161 0.0183 1.0000 10.000 1.1640 0.03040 0.02278 -0.0124 0.0196 1.0000 10.500 1.1653 0.03518 0.02794 -0.0090 0.0213 1.0000 11.000 1.1610 0.04125 0.03454 -0.0042 0.0237 1.0000 11.500 1.1749 0.04564 0.03918 -0.0021 0.0276 1.0000 12.000 1.1612 0.05640 0.05105 0.0023 0.0413 1.0000 12.500 1.1118 0.06909 0.06431 0.0017 0.0451 1.0000 13.000 1.0625 0.08180 0.07741 -0.0019 0.0460 1.0000 13.500 1.0167 0.09637 0.09228 -0.0089 0.0456 1.0000 14.000 0.9733 0.11297 0.10913 -0.0183 0.0440 1.0000 14.500 0.9292 0.13183 0.12817 -0.0286 0.0414 1.0000 15.000 0.8804 0.15567 0.15209 -0.0396 0.0394 1.0000