XFOIL Version 6.96 Calculated polar for: FX 63-100 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 ------ -------- --------- --------- -------- -------- -------- -9.000 -0.3464 0.09319 0.08964 -0.0394 1.0000 0.0178 -8.750 -0.3513 0.08989 0.08640 -0.0395 1.0000 0.0179 -8.500 -0.3583 0.08692 0.08349 -0.0389 1.0000 0.0178 -8.250 -0.3608 0.08293 0.07957 -0.0407 0.9984 0.0178 -8.000 -0.3502 0.07680 0.07345 -0.0478 0.9918 0.0177 -7.750 -0.3410 0.06970 0.06638 -0.0565 0.9816 0.0177 -7.500 -0.3268 0.05817 0.05485 -0.0733 0.9687 0.0177 -7.250 -0.2925 0.02852 0.02358 -0.1155 0.9529 0.0182 -7.000 -0.2579 0.02449 0.01889 -0.1204 0.9469 0.0200 -6.750 -0.2228 0.02346 0.01773 -0.1229 0.9432 0.0212 -6.500 -0.1883 0.02166 0.01560 -0.1254 0.9385 0.0228 -6.250 -0.1535 0.01979 0.01334 -0.1277 0.9333 0.0247 -6.000 -0.1165 0.01843 0.01174 -0.1302 0.9298 0.0273 -5.750 -0.0795 0.01774 0.01097 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0.01321 0.00601 -0.1387 0.4070 1.0000 5.000 1.1552 0.01371 0.00638 -0.1380 0.3753 1.0000 5.250 1.1779 0.01427 0.00679 -0.1372 0.3385 1.0000 5.500 1.1997 0.01490 0.00727 -0.1363 0.3000 1.0000 5.750 1.2209 0.01561 0.00779 -0.1354 0.2617 1.0000 6.000 1.2413 0.01639 0.00838 -0.1344 0.2259 1.0000 6.250 1.2622 0.01713 0.00900 -0.1334 0.1977 1.0000 6.750 1.3026 0.01872 0.01035 -0.1313 0.1420 1.0000 7.000 1.3216 0.01960 0.01109 -0.1301 0.1138 1.0000 7.250 1.3403 0.02048 0.01188 -0.1289 0.0941 1.0000 7.500 1.3590 0.02133 0.01270 -0.1276 0.0819 1.0000 7.750 1.3772 0.02219 0.01358 -0.1263 0.0730 1.0000 8.000 1.3949 0.02305 0.01447 -0.1248 0.0660 1.0000 8.250 1.4116 0.02395 0.01543 -0.1233 0.0612 1.0000 8.500 1.4283 0.02480 0.01636 -0.1217 0.0573 1.0000 8.750 1.4420 0.02574 0.01735 -0.1196 0.0540 1.0000 9.000 1.4525 0.02684 0.01849 -0.1172 0.0514 1.0000 9.250 1.4653 0.02779 0.01958 -0.1151 0.0495 1.0000 9.500 1.4769 0.02883 0.02074 -0.1129 0.0477 1.0000 9.750 1.4881 0.02992 0.02192 -0.1109 0.0457 1.0000 10.000 1.4978 0.03115 0.02320 -0.1088 0.0435 1.0000 10.250 1.5034 0.03276 0.02484 -0.1064 0.0415 1.0000 10.500 1.5146 0.03397 0.02621 -0.1047 0.0399 1.0000 10.750 1.5244 0.03534 0.02771 -0.1031 0.0381 1.0000 11.000 1.5330 0.03683 0.02931 -0.1014 0.0363 1.0000 11.250 1.5394 0.03858 0.03117 -0.0998 0.0347 1.0000 11.500 1.5413 0.04088 0.03352 -0.0981 0.0331 1.0000 11.750 1.5494 0.04263 0.03544 -0.0968 0.0318 1.0000 12.000 1.5561 0.04456 0.03754 -0.0956 0.0303 1.0000 12.250 1.5617 0.04664 0.03975 -0.0945 0.0287 1.0000 12.500 1.5654 0.04897 0.04220 -0.0936 0.0273 1.0000 12.750 1.5647 0.05190 0.04520 -0.0927 0.0261 1.0000 13.000 1.5679 0.05452 0.04801 -0.0920 0.0249 1.0000 13.250 1.5713 0.05719 0.05089 -0.0915 0.0234 1.0000 13.500 1.5728 0.06014 0.05400 -0.0912 0.0220 1.0000 13.750 1.5734 0.06326 0.05726 -0.0911 0.0209 1.0000 14.000 1.5715 0.06680 0.06092 -0.0914 0.0200 1.0000 14.250 1.5687 0.07066 0.06497 -0.0917 0.0191 1.0000 14.500 1.5672 0.07448 0.06901 -0.0922 0.0180 1.0000 14.750 1.5634 0.07875 0.07346 -0.0930 0.0172 1.0000 15.000 1.5592 0.08323 0.07810 -0.0941 0.0165 1.0000 15.250 1.5543 0.08800 0.08303 -0.0956 0.0159 1.0000 15.500 1.5481 0.09310 0.08828 -0.0974 0.0154 1.0000 15.750 1.5402 0.09868 0.09400 -0.0997 0.0150 1.0000 16.000 1.5304 0.10479 0.10023 -0.1023 0.0147 1.0000 16.250 1.5213 0.11093 0.10658 -0.1050 0.0143 1.0000 16.500 1.5112 0.11745 0.11331 -0.1081 0.0140 1.0000 16.750 1.5004 0.12427 0.12034 -0.1117 0.0136 1.0000 17.000 1.4896 0.13131 0.12756 -0.1156 0.0133 1.0000 17.250 1.4777 0.13877 0.13521 -0.1200 0.0131 1.0000 17.500 1.4650 0.14665 0.14327 -0.1249 0.0129 1.0000 17.750 1.4520 0.15487 0.15167 -0.1303 0.0127 1.0000 18.000 1.4369 0.16390 0.16088 -0.1363 0.0126 1.0000 18.250 1.4199 0.17393 0.17109 -0.1433 0.0125 1.0000