XFOIL Version 6.94 Calculated polar for: BOEING-VERTOL VR-5 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 ------- -------- --------- --------- -------- ------- ------- 1.000 0.3889 0.00947 0.00397 -0.0398 0.6342 0.9669 1.500 0.4678 0.00948 0.00384 -0.0448 0.6244 0.9880 2.000 0.5386 0.00960 0.00390 -0.0487 0.6146 1.0000 2.500 0.5909 0.00973 0.00397 -0.0487 0.6054 1.0000 3.000 0.6470 0.00998 0.00414 -0.0492 0.5963 1.0000 3.500 0.7040 0.01018 0.00438 -0.0499 0.5867 1.0000 4.500 0.8177 0.01064 0.00489 -0.0510 0.5654 1.0000 5.000 0.8741 0.01086 0.00512 -0.0513 0.5521 1.0000 5.500 0.9300 0.01102 0.00531 -0.0515 0.5351 1.0000 6.000 0.9851 0.01112 0.00547 -0.0516 0.5112 1.0000 6.500 1.0394 0.01117 0.00563 -0.0517 0.4733 1.0000 7.000 1.0915 0.01160 0.00600 -0.0516 0.4159 1.0000 7.500 1.1334 0.01318 0.00705 -0.0507 0.3076 1.0000 8.000 1.1666 0.01564 0.00892 -0.0493 0.1994 1.0000 8.500 1.1847 0.01912 0.01160 -0.0466 0.0795 1.0000 9.000 1.2049 0.02180 0.01408 -0.0440 0.0448 1.0000 9.500 1.2233 0.02467 0.01694 -0.0417 0.0314 1.0000 10.000 1.2401 0.02783 0.02020 -0.0397 0.0266 1.0000 10.500 1.2544 0.03139 0.02390 -0.0380 0.0241 1.0000 11.000 1.2674 0.03518 0.02779 -0.0368 0.0217 1.0000 11.500 1.2871 0.03842 0.03117 -0.0360 0.0188 1.0000 12.500 1.3066 0.04710 0.04021 -0.0342 0.0168 1.0000 13.000 1.3137 0.05182 0.04509 -0.0336 0.0160 1.0000 13.500 1.3168 0.05718 0.05059 -0.0333 0.0154 1.0000 14.000 1.3229 0.06243 0.05607 -0.0331 0.0148 1.0000 14.500 1.3281 0.06794 0.06182 -0.0333 0.0142 1.0000 15.000 1.3315 0.07388 0.06799 -0.0339 0.0137 1.0000 15.500 1.3335 0.08033 0.07462 -0.0353 0.0130 1.0000 16.000 1.3313 0.08759 0.08209 -0.0371 0.0121 1.0000 16.500 1.3284 0.09517 0.08997 -0.0392 0.0116 1.0000 17.000 1.3228 0.10341 0.09848 -0.0419 0.0112 1.0000 17.500 1.3154 0.11224 0.10757 -0.0452 0.0108 1.0000 18.000 1.3062 0.12163 0.11721 -0.0492 0.0105 1.0000 18.500 1.2955 0.13165 0.12748 -0.0539 0.0103 1.0000 19.000 1.2842 0.14211 0.13817 -0.0594 0.0101 1.0000 19.500 1.2717 0.15321 0.14950 -0.0658 0.0099 1.0000 20.000 1.2579 0.16506 0.16160 -0.0730 0.0099 1.0000 20.500 1.2410 0.17836 0.17517 -0.0817 0.0098 1.0000