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 = 0.200 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.500 -0.6356 0.08871 0.08528 0.0077 1.0000 0.0086 -8.250 -0.6378 0.08404 0.08065 0.0048 1.0000 0.0084 -8.000 -0.6412 0.07957 0.07623 0.0013 1.0000 0.0081 -7.750 -0.6448 0.07412 0.07078 -0.0038 1.0000 0.0081 -7.500 -0.6437 0.06874 0.06534 -0.0077 1.0000 0.0078 -7.250 -0.6404 0.06331 0.05981 -0.0106 1.0000 0.0077 -7.000 -0.6343 0.05811 0.05445 -0.0127 1.0000 0.0076 -6.750 -0.6254 0.05302 0.04917 -0.0140 1.0000 0.0074 -6.500 -0.6140 0.04803 0.04392 -0.0146 1.0000 0.0074 -6.250 -0.6000 0.04330 0.03888 -0.0147 1.0000 0.0078 -6.000 -0.5833 0.03892 0.03416 -0.0143 1.0000 0.0085 -5.750 -0.5635 0.03399 0.02868 -0.0130 1.0000 0.0096 -5.500 -0.5439 0.03001 0.02416 -0.0118 1.0000 0.0098 -5.250 -0.5232 0.02670 0.02024 -0.0106 1.0000 0.0099 -5.000 -0.5006 0.02389 0.01688 -0.0094 1.0000 0.0101 -4.750 -0.4779 0.02129 0.01394 -0.0085 1.0000 0.0104 -4.500 -0.4546 0.01948 0.01200 -0.0079 1.0000 0.0112 -4.250 -0.4302 0.01789 0.01018 -0.0071 1.0000 0.0126 -4.000 -0.4054 0.01674 0.00874 -0.0063 1.0000 0.0168 -3.750 -0.3821 0.01518 0.00714 -0.0053 1.0000 0.0193 -3.500 -0.3587 0.01401 0.00583 -0.0043 1.0000 0.0213 -3.250 -0.3364 0.01294 0.00477 -0.0035 1.0000 0.0255 -3.000 -0.3127 0.01223 0.00397 -0.0027 1.0000 0.0264 -2.750 -0.2884 0.01174 0.00336 -0.0020 1.0000 0.0281 -2.500 -0.2387 0.01136 0.00277 -0.0068 0.9296 0.0327 -2.250 -0.1992 0.01120 0.00228 -0.0089 0.8251 0.0400 -2.000 -0.1770 0.01148 0.00191 -0.0072 0.6841 0.0438 -1.750 -0.1532 0.01165 0.00158 -0.0063 0.5972 0.0455 -1.500 -0.1270 0.01161 0.00134 -0.0060 0.5668 0.0478 -1.250 -0.1011 0.01134 0.00112 -0.0057 0.5467 0.1010 -1.000 -0.0793 0.01020 0.00088 -0.0053 0.5297 0.3710 -0.750 -0.0517 0.01005 0.00078 -0.0053 0.5427 0.3903 -0.500 -0.0202 0.01109 0.00075 -0.0053 0.5263 0.1415 -0.250 0.0087 0.01130 0.00074 -0.0053 0.5410 0.0507 0.000 0.0368 0.01120 0.00072 -0.0054 0.5587 0.0466 0.250 0.0647 0.01110 0.00073 -0.0054 0.5734 0.0456 0.500 0.0922 0.01105 0.00076 -0.0054 0.5758 0.0450 0.750 0.1196 0.01103 0.00081 -0.0053 0.5727 0.0446 1.000 0.1464 0.01106 0.00086 -0.0051 0.5581 0.0443 1.250 0.1729 0.01114 0.00094 -0.0048 0.5327 0.0448 1.500 0.1986 0.01131 0.00105 -0.0044 0.4862 0.0458 1.750 0.2193 0.01179 0.00121 -0.0037 0.2934 0.2074 2.000 0.2292 0.01134 0.00163 -0.0014 0.0457 0.7323 2.250 0.2550 0.01115 0.00194 -0.0005 0.0396 0.8650 2.500 0.3123 0.01132 0.00241 -0.0068 0.0322 1.0000 2.750 0.3376 0.01158 0.00276 -0.0063 0.0288 1.0000 3.000 0.3627 0.01191 0.00323 -0.0057 0.0273 1.0000 3.250 0.3877 0.01231 0.00370 -0.0052 0.0267 1.0000 3.500 0.4124 0.01280 0.00431 -0.0046 0.0265 1.0000 3.750 0.4366 0.01343 0.00506 -0.0039 0.0264 1.0000 4.000 0.4603 0.01419 0.00588 -0.0033 0.0248 1.0000 4.250 0.4826 0.01540 0.00710 -0.0024 0.0220 1.0000 4.500 0.5045 0.01703 0.00871 -0.0016 0.0172 1.0000 4.750 0.5289 0.01845 0.01052 -0.0007 0.0142 1.0000 5.000 0.5533 0.02014 0.01246 0.0002 0.0125 1.0000 5.250 0.5763 0.02220 0.01455 0.0008 0.0115 1.0000 5.500 0.6002 0.02445 0.01742 0.0021 0.0107 1.0000 5.750 0.6236 0.02686 0.02046 0.0034 0.0083 1.0000 6.000 0.6442 0.03004 0.02411 0.0046 0.0077 1.0000 6.250 0.6626 0.03376 0.02829 0.0058 0.0074 1.0000 6.500 0.6787 0.03806 0.03305 0.0070 0.0074 1.0000 6.750 0.6923 0.04278 0.03817 0.0078 0.0075 1.0000 7.000 0.7030 0.04785 0.04360 0.0082 0.0077 1.0000 7.250 0.7108 0.05308 0.04912 0.0082 0.0079 1.0000 7.500 0.7152 0.05839 0.05467 0.0074 0.0081 1.0000 7.750 0.7165 0.06369 0.06015 0.0060 0.0084 1.0000 8.000 0.7145 0.06905 0.06563 0.0037 0.0086 1.0000 8.250 0.7086 0.07449 0.07115 0.0004 0.0088 1.0000 8.500 0.6992 0.08061 0.07725 -0.0055 0.0089 1.0000 8.750 0.6941 0.08626 0.08285 -0.0099 0.0090 1.0000