XFOIL Version 6.96 Calculated polar for: BOEING-VERTOL VR-13 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.750 -0.5086 0.10989 0.10343 0.0014 1.0000 0.2331 -8.500 -0.5011 0.10609 0.09968 0.0019 1.0000 0.2480 -8.250 -0.4996 0.10279 0.09646 0.0019 1.0000 0.2633 -8.000 -0.4733 0.09787 0.09153 0.0051 1.0000 0.2861 -7.750 -0.4729 0.09494 0.08868 0.0054 1.0000 0.3061 -7.500 -0.4850 0.09270 0.08656 0.0049 1.0000 0.3242 -7.250 -0.4631 0.08852 0.08236 0.0078 1.0000 0.3521 -7.000 -0.4528 0.08520 0.07909 0.0100 1.0000 0.3812 -6.750 -0.4557 0.08291 0.07690 0.0124 1.0000 0.4149 -6.500 -0.4218 0.07901 0.07297 0.0173 1.0000 0.4718 -6.250 -0.4221 0.07723 0.07128 0.0218 1.0000 0.5228 -6.000 -0.3734 0.07237 0.06634 0.0248 1.0000 0.5889 -5.750 -0.3452 0.06975 0.06368 0.0297 1.0000 0.6752 -5.000 -0.2999 0.04762 0.04238 0.0169 1.0000 0.5861 -4.500 -0.4454 0.04425 0.03612 -0.0105 1.0000 0.1734 -4.250 -0.4264 0.04071 0.03213 -0.0090 1.0000 0.1557 -4.000 -0.4087 0.03781 0.02892 -0.0073 1.0000 0.1501 -3.750 -0.3895 0.03540 0.02601 -0.0056 1.0000 0.1474 -3.500 -0.3673 0.03365 0.02347 -0.0038 1.0000 0.1425 -3.250 -0.3443 0.03151 0.02106 -0.0026 1.0000 0.1416 -3.000 -0.3210 0.02964 0.01897 -0.0016 1.0000 0.1452 -2.750 -0.2966 0.02815 0.01733 -0.0007 1.0000 0.1498 -2.500 -0.2700 0.02684 0.01577 0.0001 1.0000 0.1530 -2.250 -0.2421 0.02567 0.01448 0.0005 1.0000 0.1611 -2.000 -0.2113 0.02476 0.01342 0.0005 1.0000 0.1726 -1.750 -0.1830 0.02380 0.01248 0.0008 1.0000 0.1862 -1.500 -0.0637 0.01931 0.01088 -0.0127 1.0000 1.0000 -1.250 -0.0516 0.01948 0.01065 -0.0102 1.0000 1.0000 -1.000 -0.0390 0.01967 0.01054 -0.0080 1.0000 1.0000 -0.750 -0.0259 0.01989 0.01052 -0.0059 1.0000 1.0000 -0.500 -0.0120 0.02014 0.01055 -0.0041 1.0000 1.0000 -0.250 0.0024 0.02042 0.01065 -0.0024 1.0000 1.0000 0.000 0.0175 0.02074 0.01079 -0.0009 1.0000 1.0000 0.250 0.0329 0.02110 0.01101 0.0005 1.0000 1.0000 0.500 0.0487 0.02150 0.01129 0.0017 1.0000 1.0000 0.750 0.0647 0.02195 0.01163 0.0027 1.0000 1.0000 1.000 0.0808 0.02245 0.01205 0.0037 1.0000 1.0000 1.250 0.0968 0.02302 0.01256 0.0044 1.0000 1.0000 1.500 0.1428 0.02397 0.01346 -0.0007 0.9883 1.0000 1.750 0.2278 0.02509 0.01459 -0.0125 0.9576 1.0000 2.000 0.3127 0.02567 0.01527 -0.0232 0.9236 1.0000 2.250 0.3995 0.02547 0.01527 -0.0330 0.8854 1.0000 2.500 0.4767 0.02441 0.01445 -0.0390 0.8442 1.0000 2.750 0.5219 0.02316 0.01333 -0.0382 0.7967 1.0000 3.000 0.5539 0.02181 0.01201 -0.0343 0.7419 1.0000 3.250 0.5743 0.02097 0.01097 -0.0292 0.6698 1.0000 3.500 0.5945 0.02074 0.01030 -0.0250 0.5972 1.0000 3.750 0.6158 0.02117 0.01025 -0.0223 0.5403 1.0000 4.000 0.6384 0.02194 0.01066 -0.0206 0.4983 1.0000 4.250 0.6618 0.02284 0.01131 -0.0194 0.4660 1.0000 4.500 0.6858 0.02380 0.01204 -0.0183 0.4407 1.0000 4.750 0.7096 0.02485 0.01309 -0.0175 0.4193 1.0000 5.000 0.7335 0.02591 0.01408 -0.0168 0.4012 1.0000 5.250 0.7572 0.02700 0.01514 -0.0160 0.3853 1.0000 5.500 0.7807 0.02824 0.01643 -0.0153 0.3719 1.0000 5.750 0.8041 0.02961 0.01792 -0.0147 0.3608 1.0000 6.000 0.8276 0.03097 0.01929 -0.0141 0.3504 1.0000 6.250 0.8492 0.03242 0.02091 -0.0133 0.3394 1.0000 6.500 0.8693 0.03416 0.02290 -0.0125 0.3303 1.0000 6.750 0.8919 0.03583 0.02466 -0.0119 0.3229 1.0000 7.000 0.9079 0.03804 0.02726 -0.0110 0.3153 1.0000 7.250 0.9309 0.03964 0.02882 -0.0102 0.3067 1.0000 7.500 0.9405 0.04239 0.03210 -0.0090 0.2997 1.0000 7.750 0.9633 0.04412 0.03382 -0.0082 0.2916 1.0000 8.000 0.9657 0.04740 0.03764 -0.0067 0.2839 1.0000 8.250 0.9851 0.04929 0.03955 -0.0057 0.2749 1.0000 8.500 0.9803 0.05330 0.04404 -0.0042 0.2685 1.0000 8.750 0.9999 0.05530 0.04606 -0.0031 0.2600 1.0000 9.000 0.9814 0.06097 0.05217 -0.0018 0.2579 1.0000 9.250 0.9554 0.06748 0.05894 -0.0013 0.2575 1.0000 9.500 0.9227 0.07490 0.06648 -0.0019 0.2590 1.0000 9.750 0.8922 0.08253 0.07414 -0.0034 0.2618 1.0000