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 = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.500 -0.6405 0.08431 0.08275 0.0050 1.0000 0.0055 -8.250 -0.6441 0.07984 0.07830 0.0018 1.0000 0.0056 -8.000 -0.6541 0.07471 0.07320 -0.0031 1.0000 0.0056 -7.750 -0.6552 0.06895 0.06738 -0.0079 1.0000 0.0056 -7.500 -0.6534 0.06354 0.06189 -0.0111 1.0000 0.0056 -7.250 -0.6486 0.05841 0.05666 -0.0133 1.0000 0.0057 -7.000 -0.6406 0.05354 0.05166 -0.0146 1.0000 0.0060 -6.750 -0.6302 0.04878 0.04674 -0.0154 1.0000 0.0062 -6.500 -0.6162 0.04399 0.04175 -0.0154 1.0000 0.0069 -5.500 -0.5494 0.02470 0.02115 -0.0107 1.0000 0.0075 -5.250 -0.5306 0.02096 0.01702 -0.0093 1.0000 0.0073 -5.000 -0.5079 0.01890 0.01463 -0.0081 1.0000 0.0076 -4.750 -0.4830 0.01834 0.01381 -0.0072 1.0000 0.0080 -4.500 -0.4623 0.01454 0.00964 -0.0056 1.0000 0.0077 -4.250 -0.4387 0.01278 0.00764 -0.0044 1.0000 0.0078 -4.000 -0.4170 0.01091 0.00567 -0.0032 1.0000 0.0089 -3.750 -0.3926 0.01054 0.00531 -0.0027 1.0000 0.0102 -3.500 -0.3683 0.01003 0.00476 -0.0019 1.0000 0.0117 -3.250 -0.3434 0.00977 0.00446 -0.0012 1.0000 0.0131 -3.000 -0.2878 0.00879 0.00291 -0.0075 0.8016 0.0184 -2.750 -0.2665 0.00880 0.00235 -0.0061 0.6583 0.0203 -2.500 -0.2443 0.00930 0.00198 -0.0054 0.4059 0.0214 -2.250 -0.2210 0.00986 0.00152 -0.0049 0.0476 0.0242 -2.000 -0.1939 0.00972 0.00129 -0.0047 0.0433 0.0247 -1.750 -0.1667 0.00962 0.00110 -0.0046 0.0419 0.0248 -1.500 -0.1394 0.00954 0.00095 -0.0045 0.0417 0.0253 -1.250 -0.1121 0.00948 0.00084 -0.0044 0.0416 0.0249 -1.000 -0.0848 0.00943 0.00076 -0.0043 0.0416 0.0245 -0.750 -0.0574 0.00940 0.00070 -0.0042 0.0408 0.0242 -0.500 -0.0300 0.00938 0.00065 -0.0041 0.0390 0.0239 -0.250 -0.0026 0.00936 0.00062 -0.0041 0.0363 0.0237 0.000 0.0246 0.00940 0.00055 -0.0040 0.0297 0.0235 0.500 0.0794 0.00941 0.00057 -0.0038 0.0271 0.0237 0.750 0.1068 0.00943 0.00061 -0.0038 0.0270 0.0246 1.000 0.1341 0.00943 0.00069 -0.0037 0.0271 0.0349 1.250 0.1615 0.00946 0.00075 -0.0036 0.0267 0.0392 1.500 0.1888 0.00950 0.00082 -0.0035 0.0261 0.0404 1.750 0.2161 0.00955 0.00091 -0.0034 0.0258 0.0420 2.000 0.2434 0.00963 0.00106 -0.0033 0.0253 0.0507 2.250 0.2645 0.00824 0.00120 -0.0028 0.0245 0.5968 2.500 0.2879 0.00795 0.00150 -0.0019 0.0232 0.7424 2.750 0.3305 0.00770 0.00208 -0.0048 0.0229 0.9894 3.000 0.3683 0.00814 0.00259 -0.0071 0.0225 1.0000 3.250 0.3925 0.00861 0.00315 -0.0063 0.0222 1.0000 3.500 0.4161 0.00915 0.00371 -0.0056 0.0193 1.0000 3.750 0.4373 0.01031 0.00489 -0.0045 0.0147 1.0000 4.000 0.4628 0.01052 0.00513 -0.0040 0.0129 1.0000 4.250 0.4874 0.01103 0.00566 -0.0034 0.0108 1.0000 4.500 0.5110 0.01180 0.00640 -0.0029 0.0096 1.0000 4.750 0.5342 0.01325 0.00796 -0.0019 0.0086 1.0000 5.000 0.5600 0.01374 0.00863 -0.0013 0.0076 1.0000 5.250 0.5844 0.01509 0.01017 -0.0004 0.0070 1.0000 5.500 0.6089 0.01606 0.01129 0.0003 0.0063 1.0000 5.750 0.6332 0.01664 0.01198 0.0006 0.0055 1.0000 6.000 0.6281 0.01189 0.00832 0.0040 0.0051 1.0000 6.250 0.6494 0.01352 0.01014 0.0047 0.0053 1.0000 8.500 0.6230 0.06869 0.06725 -0.0036 0.0080 1.0000 8.750 0.6139 0.07469 0.07323 -0.0068 0.0082 1.0000 9.000 0.6074 0.08080 0.07933 -0.0098 0.0080 1.0000 9.250 0.6028 0.08658 0.08510 -0.0123 0.0079 1.0000