XFOIL Version 6.96 Calculated polar for: FX 38-153 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -12.500 -0.1825 0.08507 0.08327 -0.1164 0.9695 0.0128 -12.250 -0.1924 0.07868 0.07690 -0.1188 0.9685 0.0130 -12.000 -0.2117 0.06935 0.06758 -0.1229 0.9675 0.0130 -11.750 -0.2815 0.05076 0.04880 -0.1329 0.9659 0.0123 -11.500 -0.3156 0.04385 0.04175 -0.1347 0.9634 0.0122 -11.250 -0.3378 0.03939 0.03716 -0.1347 0.9608 0.0122 -11.000 -0.3570 0.03562 0.03326 -0.1339 0.9584 0.0122 -10.750 -0.3717 0.03267 0.03018 -0.1325 0.9558 0.0123 -10.500 -0.3837 0.03025 0.02764 -0.1306 0.9529 0.0124 -10.250 -0.3932 0.02807 0.02535 -0.1283 0.9497 0.0127 -10.000 -0.4010 0.02599 0.02312 -0.1258 0.9461 0.0130 -9.750 -0.4521 0.03690 0.03369 -0.1233 0.9444 0.0127 -9.500 -0.4522 0.03601 0.03227 -0.1196 0.9414 0.0144 -9.250 -0.4506 0.03471 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0.9093 0.01408 0.00711 -0.0831 0.0707 0.8077 5.750 0.9298 0.01455 0.00748 -0.0821 0.0548 0.8089 6.000 0.9527 0.01488 0.00778 -0.0816 0.0485 0.8101 6.250 0.9746 0.01527 0.00813 -0.0810 0.0362 0.8112 6.500 0.9966 0.01566 0.00848 -0.0803 0.0306 0.8123 6.750 1.0195 0.01599 0.00880 -0.0799 0.0289 0.8133 7.000 1.0424 0.01631 0.00913 -0.0794 0.0274 0.8143 7.250 1.0652 0.01663 0.00947 -0.0789 0.0261 0.8153 7.500 1.0880 0.01695 0.00982 -0.0785 0.0249 0.8161 8.000 1.1292 0.01781 0.01066 -0.0769 0.0163 0.8183 8.250 1.1495 0.01822 0.01109 -0.0760 0.0158 0.8192 8.500 1.1696 0.01865 0.01156 -0.0751 0.0155 0.8200 8.750 1.1892 0.01912 0.01205 -0.0742 0.0150 0.8209 9.000 1.2084 0.01961 0.01258 -0.0732 0.0146 0.8218 9.250 1.2270 0.02014 0.01313 -0.0721 0.0141 0.8228 9.500 1.2454 0.02068 0.01372 -0.0710 0.0136 0.8237 9.750 1.2635 0.02126 0.01434 -0.0699 0.0133 0.8247 10.000 1.2812 0.02186 0.01499 -0.0688 0.0130 0.8257 10.250 1.2985 0.02248 0.01566 -0.0676 0.0128 0.8267 10.500 1.3154 0.02312 0.01636 -0.0664 0.0126 0.8278 10.750 1.3318 0.02381 0.01711 -0.0652 0.0123 0.8289 11.000 1.3472 0.02457 0.01791 -0.0639 0.0121 0.8299 11.250 1.3618 0.02538 0.01878 -0.0625 0.0117 0.8308 11.500 1.3754 0.02627 0.01973 -0.0610 0.0113 0.8315 11.750 1.3872 0.02727 0.02080 -0.0594 0.0110 0.8322 12.000 1.3978 0.02838 0.02198 -0.0577 0.0106 0.8334 12.250 1.4055 0.02967 0.02336 -0.0556 0.0102 0.8344 12.500 1.4189 0.03055 0.02431 -0.0543 0.0099 0.8353 12.750 1.4343 0.03129 0.02511 -0.0533 0.0097 0.8362 13.000 1.4495 0.03205 0.02592 -0.0523 0.0092 0.8371 13.250 1.4631 0.03296 0.02689 -0.0512 0.0087 0.8380 13.500 1.4768 0.03388 0.02784 -0.0502 0.0080 0.8390 13.750 1.4842 0.03531 0.02934 -0.0486 0.0074 0.8400 14.000 1.5049 0.03567 0.02972 -0.0484 0.0068 0.8412 14.250 1.5203 0.03648 0.03058 -0.0476 0.0057 0.8424 14.500 1.5328 0.03755 0.03162 -0.0467 0.0035 0.8435 14.750 1.5398 0.03912 0.03320 -0.0454 0.0031 0.8446 15.000 1.5453 0.04088 0.03502 -0.0440 0.0027 0.8455 15.250 1.5510 0.04269 0.03689 -0.0428 0.0025 0.8463 15.500 1.5554 0.04465 0.03894 -0.0416 0.0025 0.8470 15.750 1.5587 0.04679 0.04118 -0.0406 0.0024 0.8480 16.000 1.5606 0.04913 0.04363 -0.0396 0.0023 0.8491 16.250 1.5613 0.05167 0.04630 -0.0387 0.0023 0.8502 16.500 1.5606 0.05445 0.04919 -0.0380 0.0022 0.8513 16.750 1.5586 0.05746 0.05233 -0.0374 0.0022 0.8524 17.000 1.5562 0.06062 0.05561 -0.0371 0.0022 0.8534 17.250 1.5519 0.06414 0.05926 -0.0370 0.0021 0.8544 17.500 1.5462 0.06799 0.06323 -0.0372 0.0021 0.8554 17.750 1.5393 0.07212 0.06750 -0.0376 0.0021 0.8563 18.000 1.5302 0.07666 0.07217 -0.0384 0.0021 0.8572 18.250 1.5190 0.08165 0.07729 -0.0395 0.0020 0.8581 18.500 1.5071 0.08689 0.08267 -0.0409 0.0020 0.8589 18.750 1.4939 0.09248 0.08840 -0.0427 0.0020 0.8596 19.000 1.4780 0.09862 0.09467 -0.0449 0.0020 0.8603 19.250 1.4614 0.10498 0.10118 -0.0474 0.0020 0.8609