XFOIL Version 6.94 Calculated polar for: A18 (original) 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.6453 0.00801 0.00244 -0.1261 0.6911 0.5614 0.500 0.6991 0.00770 0.00263 -0.1254 0.6613 0.7603 1.000 0.7474 0.00752 0.00263 -0.1230 0.6273 1.0000 1.500 0.8025 0.00786 0.00278 -0.1227 0.5942 1.0000 2.000 0.8563 0.00833 0.00301 -0.1221 0.5535 1.0000 2.500 0.9105 0.00879 0.00331 -0.1215 0.5201 1.0000 3.000 0.9646 0.00919 0.00358 -0.1211 0.4843 1.0000 3.500 1.0160 0.00977 0.00390 -0.1202 0.4203 1.0000 4.000 1.0657 0.01061 0.00449 -0.1191 0.3720 1.0000 4.500 1.1120 0.01178 0.00513 -0.1176 0.2560 1.0000 5.000 1.1566 0.01311 0.00610 -0.1160 0.1963 1.0000 5.500 1.2055 0.01392 0.00684 -0.1148 0.1767 1.0000 6.000 1.2535 0.01475 0.00760 -0.1136 0.1537 1.0000 6.500 1.3005 0.01561 0.00838 -0.1122 0.1260 1.0000 7.000 1.3440 0.01672 0.00931 -0.1103 0.0953 1.0000 7.500 1.3841 0.01802 0.01051 -0.1079 0.0755 1.0000 8.000 1.4216 0.01938 0.01183 -0.1051 0.0650 1.0000 8.500 1.4575 0.02069 0.01321 -0.1020 0.0570 1.0000 9.000 1.4879 0.02214 0.01471 -0.0980 0.0505 1.0000 9.500 1.5105 0.02383 0.01645 -0.0929 0.0449 1.0000 10.000 1.5316 0.02579 0.01850 -0.0881 0.0402 1.0000 10.500 1.5521 0.02790 0.02073 -0.0837 0.0361 1.0000 11.000 1.5725 0.03017 0.02316 -0.0798 0.0328 1.0000 11.500 1.5807 0.03361 0.02673 -0.0752 0.0311 1.0000 12.000 1.5975 0.03656 0.02985 -0.0722 0.0288 1.0000 12.500 1.6001 0.04099 0.03439 -0.0689 0.0278 1.0000 13.000 1.6106 0.04508 0.03876 -0.0667 0.0261 1.0000 13.500 1.6120 0.05025 0.04409 -0.0649 0.0251 1.0000 14.000 1.6095 0.05622 0.05026 -0.0637 0.0239 1.0000 14.500 1.6033 0.06300 0.05731 -0.0632 0.0227 1.0000 15.000 1.5998 0.06980 0.06436 -0.0634 0.0225 1.0000 15.500 1.5948 0.07718 0.07189 -0.0647 0.0217 1.0000 16.000 1.5861 0.08514 0.08008 -0.0658 0.0217 1.0000 17.000 1.5584 0.10238 0.09770 -0.0684 0.0210 1.0000