XFOIL Version 6.96 Calculated polar for: Dayton-Wright T-1 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 ------ -------- --------- --------- -------- -------- -------- -17.500 -0.3829 0.16447 0.16260 -0.0306 1.0000 0.0212 -17.250 -0.3755 0.16160 0.15973 -0.0319 1.0000 0.0219 -15.750 -0.8818 0.03810 0.03526 -0.1106 0.9974 0.0249 -15.500 -0.8727 0.03493 0.03192 -0.1141 0.9942 0.0251 -15.250 -0.8603 0.03259 0.02942 -0.1156 0.9900 0.0253 -15.000 -0.8414 0.03066 0.02736 -0.1170 0.9869 0.0256 -14.750 -0.8195 0.02890 0.02546 -0.1184 0.9848 0.0258 -14.500 -0.7968 0.02731 0.02373 -0.1195 0.9828 0.0261 -14.250 -0.7785 0.02597 0.02226 -0.1192 0.9780 0.0264 -14.000 -0.7579 0.02476 0.02092 -0.1190 0.9735 0.0266 -13.750 -0.7353 0.02379 0.01982 -0.1189 0.9698 0.0270 -13.500 -0.7166 0.02297 0.01888 -0.1177 0.9639 0.0272 -13.250 -0.6966 0.02216 0.01795 -0.1166 0.9584 0.0274 -13.000 -0.6752 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0.00846 0.00333 -0.1013 0.3645 0.9998 2.250 0.9519 0.00867 0.00347 -0.1003 0.3504 1.0000 2.500 0.9710 0.00889 0.00361 -0.0987 0.3379 1.0000 2.750 0.9903 0.00910 0.00376 -0.0971 0.3260 1.0000 3.000 1.0099 0.00929 0.00390 -0.0956 0.3152 1.0000 3.250 1.0288 0.00951 0.00406 -0.0940 0.3046 1.0000 3.500 1.0484 0.00971 0.00422 -0.0926 0.2958 1.0000 3.750 1.0679 0.00991 0.00438 -0.0911 0.2886 1.0000 4.000 1.0883 0.01008 0.00454 -0.0898 0.2823 1.0000 4.250 1.1068 0.01033 0.00473 -0.0882 0.2747 1.0000 4.500 1.1277 0.01048 0.00489 -0.0870 0.2697 1.0000 4.750 1.1478 0.01068 0.00507 -0.0857 0.2647 1.0000 5.000 1.1671 0.01093 0.00529 -0.0843 0.2594 1.0000 5.250 1.1884 0.01110 0.00547 -0.0833 0.2562 1.0000 5.500 1.2086 0.01128 0.00566 -0.0821 0.2519 1.0000 5.750 1.2271 0.01154 0.00589 -0.0806 0.2464 1.0000 6.000 1.2465 0.01178 0.00613 -0.0793 0.2415 1.0000 6.250 1.2678 0.01198 0.00634 -0.0785 0.2378 1.0000 6.500 1.2880 0.01224 0.00659 -0.0774 0.2330 1.0000 6.750 1.3066 0.01258 0.00691 -0.0762 0.2270 1.0000 7.000 1.3288 0.01278 0.00714 -0.0755 0.2237 1.0000 7.250 1.3496 0.01305 0.00742 -0.0747 0.2196 1.0000 7.500 1.3690 0.01339 0.00776 -0.0737 0.2146 1.0000 7.750 1.3892 0.01370 0.00808 -0.0729 0.2103 1.0000 8.000 1.4101 0.01398 0.00838 -0.0722 0.2056 1.0000 8.250 1.4283 0.01441 0.00878 -0.0711 0.1985 1.0000 8.500 1.4482 0.01475 0.00914 -0.0703 0.1919 1.0000 8.750 1.4653 0.01526 0.00960 -0.0692 0.1827 1.0000 9.000 1.4823 0.01577 0.01008 -0.0681 0.1704 1.0000 9.250 1.4925 0.01667 0.01082 -0.0660 0.1449 1.0000 9.500 1.4940 0.01811 0.01204 -0.0629 0.1149 1.0000 9.750 1.5018 0.01921 0.01307 -0.0607 0.1003 1.0000 10.000 1.5117 0.02022 0.01404 -0.0589 0.0906 1.0000 10.250 1.5218 0.02125 0.01503 -0.0573 0.0823 1.0000 10.500 1.5327 0.02224 0.01601 -0.0558 0.0746 1.0000 10.750 1.5405 0.02346 0.01719 -0.0541 0.0642 1.0000 11.000 1.5394 0.02534 0.01893 -0.0515 0.0448 1.0000 11.250 1.5328 0.02769 0.02117 -0.0486 0.0278 1.0000 11.500 1.5350 0.02951 0.02298 -0.0468 0.0226 1.0000 11.750 1.5399 0.03116 0.02467 -0.0454 0.0203 1.0000 12.000 1.5451 0.03284 0.02640 -0.0441 0.0190 1.0000 12.250 1.5479 0.03480 0.02841 -0.0429 0.0178 1.0000 12.500 1.5546 0.03647 0.03015 -0.0420 0.0172 1.0000 12.750 1.5594 0.03837 0.03212 -0.0412 0.0165 1.0000 13.000 1.5623 0.04052 0.03433 -0.0404 0.0159 1.0000 13.250 1.5631 0.04297 0.03685 -0.0396 0.0154 1.0000 13.500 1.5605 0.04587 0.03984 -0.0389 0.0149 1.0000 13.750 1.5619 0.04842 0.04247 -0.0386 0.0146 1.0000 14.000 1.5630 0.05108 0.04521 -0.0383 0.0143 1.0000 14.250 1.5626 0.05399 0.04821 -0.0382 0.0141 1.0000 14.500 1.5606 0.05715 0.05146 -0.0382 0.0138 1.0000 14.750 1.5571 0.06057 0.05496 -0.0383 0.0135 1.0000 15.000 1.5519 0.06427 0.05876 -0.0386 0.0133 1.0000 15.250 1.5449 0.06829 0.06287 -0.0390 0.0131 1.0000 15.500 1.5361 0.07264 0.06732 -0.0395 0.0129 1.0000 15.750 1.5245 0.07744 0.07223 -0.0402 0.0128 1.0000 16.000 1.5111 0.08260 0.07750 -0.0411 0.0126 1.0000 16.250 1.4955 0.08819 0.08320 -0.0423 0.0125 1.0000 16.500 1.4781 0.09414 0.08928 -0.0436 0.0123 1.0000