XFOIL Version 6.96 Calculated polar for: BOEING-VERTOL VR-11X 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 -0.4148 0.10834 0.10190 -0.0183 1.0000 0.2669 -8.500 -0.4306 0.10735 0.10101 -0.0162 1.0000 0.2788 -8.250 -0.4209 0.10398 0.09766 -0.0138 1.0000 0.2932 -8.000 -0.4190 0.10124 0.09496 -0.0114 1.0000 0.3072 -7.750 -0.4222 0.09890 0.09269 -0.0088 1.0000 0.3223 -7.500 -0.4263 0.09670 0.09056 -0.0060 1.0000 0.3394 -7.250 -0.4337 0.09477 0.08870 -0.0028 1.0000 0.3584 -7.000 -0.4508 0.09341 0.08745 0.0010 1.0000 0.3774 -6.750 -0.4312 0.08997 0.08402 0.0040 1.0000 0.4064 -6.500 -0.4255 0.08745 0.08156 0.0076 1.0000 0.4367 -6.250 -0.4593 0.08731 0.08156 0.0139 1.0000 0.4613 -6.000 -0.4017 0.08236 0.07653 0.0154 1.0000 0.5175 -5.750 -0.3816 0.07984 0.07403 0.0194 1.0000 0.5733 -4.500 -0.3772 0.06736 0.06184 0.0343 1.0000 0.6757 -4.000 -0.4792 0.04726 0.03895 -0.0052 1.0000 0.1641 -3.750 -0.4558 0.04363 0.03489 -0.0042 1.0000 0.1436 -3.500 -0.4318 0.04103 0.03161 -0.0028 1.0000 0.1304 -3.250 -0.4084 0.03857 0.02879 -0.0017 1.0000 0.1257 -3.000 -0.3830 0.03653 0.02617 -0.0005 1.0000 0.1209 -2.750 -0.3569 0.03532 0.02436 0.0010 1.0000 0.1174 -2.500 -0.3304 0.03337 0.02221 0.0017 1.0000 0.1161 -2.250 -0.3034 0.03185 0.02044 0.0025 1.0000 0.1156 -2.000 -0.2759 0.03067 0.01899 0.0033 1.0000 0.1172 -1.750 -0.2468 0.02925 0.01764 0.0034 1.0000 0.1237 -1.500 -0.2157 0.02838 0.01664 0.0035 1.0000 0.1317 -1.250 -0.1876 0.02745 0.01572 0.0039 1.0000 0.1434 -1.000 -0.0801 0.02339 0.01467 -0.0085 1.0000 1.0000 -0.750 -0.0628 0.02355 0.01436 -0.0066 1.0000 1.0000 -0.500 -0.0457 0.02374 0.01423 -0.0050 1.0000 1.0000 -0.250 -0.0286 0.02398 0.01420 -0.0035 1.0000 1.0000 0.000 -0.0113 0.02426 0.01426 -0.0022 1.0000 1.0000 0.250 0.0060 0.02458 0.01440 -0.0009 1.0000 1.0000 0.500 0.0232 0.02495 0.01461 0.0003 1.0000 1.0000 0.750 0.0403 0.02537 0.01490 0.0014 1.0000 1.0000 1.000 0.0571 0.02585 0.01527 0.0024 1.0000 1.0000 1.250 0.0736 0.02640 0.01573 0.0034 1.0000 1.0000 1.500 0.0897 0.02703 0.01628 0.0042 1.0000 1.0000 1.750 0.1053 0.02774 0.01694 0.0050 1.0000 1.0000 2.000 0.1202 0.02855 0.01772 0.0056 1.0000 1.0000 2.250 0.2245 0.03072 0.01987 -0.0103 0.9605 1.0000 2.500 0.3062 0.03172 0.02092 -0.0205 0.9231 1.0000 2.750 0.3719 0.03198 0.02128 -0.0270 0.8891 1.0000 3.000 0.4402 0.03153 0.02101 -0.0329 0.8544 1.0000 3.250 0.5053 0.03022 0.01991 -0.0371 0.8151 1.0000 3.500 0.5554 0.02807 0.01799 -0.0375 0.7656 1.0000 3.750 0.5969 0.02508 0.01515 -0.0350 0.6944 1.0000 4.000 0.6802 0.02232 0.01131 -0.0372 0.5550 1.0000 4.250 0.7059 0.02346 0.01185 -0.0360 0.4940 1.0000 4.500 0.7324 0.02446 0.01253 -0.0353 0.4588 1.0000 4.750 0.7607 0.02542 0.01326 -0.0351 0.4340 1.0000 5.000 0.7872 0.02639 0.01416 -0.0347 0.4145 1.0000 5.250 0.8146 0.02743 0.01513 -0.0346 0.4001 1.0000 5.500 0.8411 0.02850 0.01615 -0.0343 0.3876 1.0000 5.750 0.8640 0.02965 0.01745 -0.0336 0.3759 1.0000 6.000 0.8879 0.03088 0.01873 -0.0331 0.3660 1.0000 6.250 0.9117 0.03216 0.02009 -0.0326 0.3575 1.0000 6.500 0.9328 0.03363 0.02173 -0.0318 0.3494 1.0000 6.750 0.9552 0.03500 0.02319 -0.0310 0.3409 1.0000 7.000 0.9733 0.03671 0.02515 -0.0299 0.3338 1.0000 7.250 0.9921 0.03838 0.02700 -0.0289 0.3266 1.0000 7.500 1.0105 0.04020 0.02897 -0.0278 0.3197 1.0000 7.750 1.0220 0.04225 0.03136 -0.0260 0.3125 1.0000 8.000 1.0469 0.04398 0.03300 -0.0256 0.3056 1.0000 8.250 1.0467 0.04671 0.03627 -0.0227 0.2999 1.0000 8.500 1.0578 0.04894 0.03870 -0.0209 0.2932 1.0000 8.750 1.0721 0.05139 0.04129 -0.0196 0.2875 1.0000 9.000 1.0604 0.05513 0.04547 -0.0163 0.2837 1.0000 9.250 1.0533 0.05867 0.04928 -0.0135 0.2793 1.0000 9.500 1.0808 0.06050 0.05108 -0.0132 0.2723 1.0000 9.750 1.0547 0.06558 0.05648 -0.0097 0.2715 1.0000 10.000 1.0241 0.07127 0.06237 -0.0068 0.2717 1.0000 10.250 0.9932 0.07733 0.06851 -0.0047 0.2725 1.0000