XFOIL Version 6.94 Calculated polar for: BOEING-VERTOL VR-8 AIRFOIL WITH 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 0.0786 0.01190 0.00146 -0.0105 0.0342 0.0375 0.500 0.1376 0.01192 0.00135 -0.0113 0.0353 0.0373 1.000 0.1965 0.01196 0.00134 -0.0121 0.0384 0.0373 1.500 0.2530 0.01097 0.00118 -0.0115 0.2443 0.0377 2.000 0.3107 0.01031 0.00123 -0.0114 0.3666 0.0425 2.500 0.3749 0.00997 0.00209 -0.0162 0.0694 0.7812 3.000 0.4133 0.00980 0.00245 -0.0114 0.0460 1.0000 3.500 0.4729 0.01032 0.00297 -0.0126 0.0328 1.0000 4.500 0.5909 0.01159 0.00424 -0.0148 0.0238 1.0000 5.000 0.6468 0.01388 0.00630 -0.0166 0.0171 1.0000 5.500 0.7026 0.01502 0.00738 -0.0174 0.0147 1.0000 6.000 0.7599 0.01541 0.00791 -0.0181 0.0130 1.0000 6.500 0.8186 0.01567 0.00894 -0.0184 0.0054 1.0000 7.000 0.8729 0.01716 0.01062 -0.0188 0.0042 1.0000 7.500 0.9233 0.02002 0.01389 -0.0186 0.0041 1.0000 8.000 0.9677 0.02530 0.01993 -0.0176 0.0041 1.0000 8.500 0.9985 0.03388 0.02955 -0.0161 0.0043 1.0000 9.000 1.0229 0.04101 0.03742 -0.0150 0.0046 1.0000 9.500 1.0113 0.05331 0.05052 -0.0138 0.0051 1.0000 10.500 0.9345 0.08045 0.07836 -0.0298 0.0053 1.0000 11.000 0.8979 0.10582 0.10381 -0.0472 0.0053 1.0000 11.500 0.8708 0.12498 0.12290 -0.0569 0.0058 1.0000 12.000 0.8584 0.13983 0.13768 -0.0638 0.0064 1.0000 12.500 0.8555 0.15210 0.14992 -0.0694 0.0072 1.0000 13.000 0.8585 0.16237 0.16017 -0.0741 0.0081 1.0000 13.500 0.7008 0.16268 0.16059 -0.0542 0.0073 1.0000 14.000 0.6931 0.16857 0.16647 -0.0561 0.0081 1.0000