XFOIL Version 6.94 Calculated polar for: BOEING 106 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3640 0.00891 0.00380 -0.0657 0.7033 0.9654 0.500 0.4527 0.00907 0.00374 -0.0724 0.6814 0.9850 1.000 0.5479 0.00910 0.00358 -0.0809 0.6581 1.0000 1.500 0.5961 0.00918 0.00351 -0.0797 0.6361 1.0000 2.000 0.6436 0.00934 0.00353 -0.0783 0.6142 1.0000 2.500 0.6909 0.00955 0.00364 -0.0768 0.5929 1.0000 3.000 0.7377 0.00982 0.00380 -0.0751 0.5716 1.0000 3.500 0.7837 0.01013 0.00401 -0.0733 0.5493 1.0000 4.000 0.8288 0.01048 0.00426 -0.0712 0.5253 1.0000 4.500 0.8732 0.01084 0.00458 -0.0690 0.5013 1.0000 5.000 0.9111 0.01121 0.00476 -0.0656 0.4554 1.0000 5.500 0.9487 0.01172 0.00506 -0.0622 0.4092 1.0000 6.000 0.9879 0.01231 0.00549 -0.0592 0.3704 1.0000 6.500 1.0195 0.01333 0.00615 -0.0552 0.3021 1.0000 7.000 1.0539 0.01435 0.00688 -0.0518 0.2399 1.0000 7.500 1.0768 0.01612 0.00813 -0.0470 0.1550 1.0000 8.000 1.0928 0.01812 0.00964 -0.0413 0.0851 1.0000 8.500 1.1133 0.01991 0.01115 -0.0366 0.0465 1.0000 9.000 1.1239 0.02252 0.01350 -0.0314 0.0058 1.0000 9.500 1.1517 0.02418 0.01530 -0.0286 0.0050 1.0000 10.000 1.1783 0.02604 0.01730 -0.0260 0.0044 1.0000 10.500 1.2031 0.02815 0.01958 -0.0236 0.0042 1.0000 11.000 1.2255 0.03058 0.02219 -0.0214 0.0040 1.0000 11.500 1.2448 0.03340 0.02522 -0.0194 0.0040 1.0000 12.000 1.2606 0.03668 0.02876 -0.0176 0.0039 1.0000 12.500 1.2730 0.04046 0.03277 -0.0161 0.0039 1.0000 13.000 1.2817 0.04479 0.03734 -0.0150 0.0039 1.0000 13.500 1.2871 0.04970 0.04251 -0.0142 0.0039 1.0000 14.000 1.2894 0.05526 0.04832 -0.0140 0.0039 1.0000 14.500 1.2882 0.06155 0.05489 -0.0143 0.0040 1.0000 15.000 1.2835 0.06864 0.06224 -0.0153 0.0040 1.0000 15.500 1.2749 0.07656 0.07046 -0.0169 0.0040 1.0000 16.000 1.2631 0.08526 0.07943 -0.0191 0.0041 1.0000 16.500 1.2485 0.09473 0.08918 -0.0219 0.0041 1.0000 17.000 1.2321 0.10483 0.09954 -0.0254 0.0042 1.0000 17.500 1.2148 0.11541 0.11038 -0.0296 0.0042 1.0000 18.000 1.1980 0.12626 0.12148 -0.0343 0.0043 1.0000 18.500 1.1823 0.13726 0.13271 -0.0396 0.0043 1.0000 19.000 1.1684 0.14825 0.14393 -0.0453 0.0044 1.0000 19.500 1.1570 0.15911 0.15501 -0.0514 0.0045 1.0000 20.000 1.1478 0.16985 0.16595 -0.0578 0.0045 1.0000 20.500 1.1403 0.18056 0.17686 -0.0645 0.0046 1.0000 21.000 1.1330 0.19170 0.18819 -0.0718 0.0047 1.0000 21.500 1.1236 0.20391 0.20062 -0.0799 0.0049 1.0000 22.000 1.1094 0.21854 0.21546 -0.0896 0.0050 1.0000 22.500 1.0777 0.24229 0.23942 -0.1040 0.0053 1.0000