XFOIL Version 6.96 Calculated polar for: BOEING-VERTOL VR-9 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.750 -0.6394 0.08983 0.08762 0.0092 1.0000 0.0036 -8.500 -0.6414 0.08521 0.08303 0.0065 1.0000 0.0036 -8.250 -0.6456 0.08039 0.07824 0.0032 1.0000 0.0035 -8.000 -0.6521 0.07583 0.07371 -0.0010 1.0000 0.0034 -7.750 -0.6541 0.06985 0.06769 -0.0064 1.0000 0.0034 -7.500 -0.6526 0.06426 0.06202 -0.0099 1.0000 0.0033 -7.250 -0.6488 0.05858 0.05622 -0.0125 1.0000 0.0032 -7.000 -0.6418 0.05312 0.05059 -0.0141 1.0000 0.0031 -6.750 -0.6323 0.04768 0.04494 -0.0149 1.0000 0.0030 -6.500 -0.6204 0.04219 0.03917 -0.0149 1.0000 0.0029 -6.250 -0.6064 0.03672 0.03337 -0.0144 1.0000 0.0029 -6.000 -0.5904 0.03148 0.02773 -0.0132 1.0000 0.0029 -5.750 -0.5727 0.02626 0.02204 -0.0116 1.0000 0.0030 -5.500 -0.5528 0.02178 0.01707 -0.0099 1.0000 0.0034 -5.250 -0.5309 0.01862 0.01345 -0.0085 1.0000 0.0039 -4.750 -0.4590 0.01464 0.00863 -0.0121 0.8973 0.0056 -4.500 -0.4385 0.01377 0.00723 -0.0102 0.7950 0.0060 -4.250 -0.4170 0.01320 0.00614 -0.0087 0.6963 0.0069 -4.000 -0.3934 0.01262 0.00510 -0.0077 0.6182 0.0090 -3.750 -0.3728 0.01370 0.00460 -0.0070 0.0458 0.0108 -3.500 -0.3461 0.01356 0.00437 -0.0069 0.0378 0.0142 -3.250 -0.3184 0.01362 0.00424 -0.0067 0.0355 0.0155 -3.000 -0.2955 0.01239 0.00304 -0.0062 0.0349 0.0179 -2.750 -0.2698 0.01190 0.00243 -0.0058 0.0331 0.0191 -2.500 -0.2435 0.01154 0.00195 -0.0056 0.0328 0.0203 -2.250 -0.2169 0.01131 0.00160 -0.0054 0.0325 0.0211 -2.000 -0.1900 0.01117 0.00134 -0.0053 0.0322 0.0219 -1.750 -0.1630 0.01107 0.00114 -0.0051 0.0299 0.0224 -1.500 -0.1358 0.01098 0.00092 -0.0050 0.0266 0.0238 -1.250 -0.1089 0.01106 0.00088 -0.0048 0.0265 0.0237 0.500 0.0812 0.01093 0.00047 -0.0043 0.0265 0.0229 1.000 0.1354 0.01095 0.00063 -0.0042 0.0228 0.0229 1.250 0.1626 0.01095 0.00071 -0.0041 0.0223 0.0230 1.500 0.1897 0.01099 0.00082 -0.0040 0.0214 0.0231 1.750 0.2166 0.01111 0.00103 -0.0039 0.0207 0.0233 2.250 0.2702 0.01143 0.00163 -0.0035 0.0192 0.0319 2.500 0.2968 0.01166 0.00194 -0.0033 0.0189 0.0326 2.750 0.3232 0.01191 0.00228 -0.0032 0.0188 0.0348 3.000 0.3496 0.01218 0.00270 -0.0029 0.0188 0.0362 3.250 0.3759 0.01248 0.00309 -0.0027 0.0190 0.0415 3.500 0.4020 0.01283 0.00357 -0.0024 0.0193 0.0456 3.750 0.4263 0.01356 0.00432 -0.0020 0.0177 0.0472 4.500 0.4871 0.01275 0.00584 0.0016 0.0113 0.8828 4.750 0.5366 0.01314 0.00680 -0.0028 0.0074 0.9999 5.000 0.5626 0.01319 0.00685 -0.0027 0.0059 1.0000 5.250 0.5866 0.01411 0.00793 -0.0019 0.0050 1.0000 5.500 0.6098 0.01567 0.00974 -0.0009 0.0043 1.0000 5.750 0.6327 0.01783 0.01225 0.0003 0.0037 1.0000 6.000 0.6541 0.02100 0.01590 0.0017 0.0034 1.0000 6.250 0.6730 0.02537 0.02083 0.0035 0.0033 1.0000 6.500 0.6892 0.03046 0.02644 0.0053 0.0034 1.0000 6.750 0.7025 0.03617 0.03263 0.0070 0.0035 1.0000 7.000 0.7134 0.04197 0.03882 0.0081 0.0037 1.0000 7.250 0.7218 0.04763 0.04478 0.0085 0.0039 1.0000 7.500 0.7273 0.05323 0.05063 0.0084 0.0040 1.0000 7.750 0.7297 0.05871 0.05630 0.0075 0.0041 1.0000 8.000 0.7289 0.06416 0.06189 0.0058 0.0042 1.0000 8.250 0.7244 0.06976 0.06759 0.0030 0.0043 1.0000 8.500 0.7127 0.07565 0.07353 -0.0018 0.0044 1.0000 8.750 0.7037 0.08192 0.07976 -0.0078 0.0044 1.0000 9.000 0.6989 0.08737 0.08518 -0.0115 0.0045 1.0000