XFOIL Version 6.96 Calculated polar for: Smoothed ATR airfoil coordinates obtained using 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.250 -0.6967 0.04843 0.04363 -0.0382 1.0000 0.0471 -9.000 -0.7067 0.04574 0.04081 -0.0339 1.0000 0.0469 -8.750 -0.7170 0.04310 0.03798 -0.0295 1.0000 0.0468 -8.500 -0.7272 0.04062 0.03516 -0.0249 1.0000 0.0470 -8.000 -0.7275 0.03565 0.02978 -0.0184 0.9997 0.0479 -7.750 -0.6890 0.03412 0.02826 -0.0215 0.9946 0.0491 -7.500 -0.6543 0.03149 0.02531 -0.0239 0.9878 0.0500 -7.250 -0.6138 0.02893 0.02234 -0.0271 0.9820 0.0514 -7.000 -0.5790 0.02680 0.01981 -0.0287 0.9739 0.0530 -6.750 -0.5363 0.02485 0.01756 -0.0319 0.9693 0.0556 -6.500 -0.5006 0.02393 0.01667 -0.0335 0.9611 0.0583 -6.250 -0.4551 0.02272 0.01521 -0.0368 0.9564 0.0620 -6.000 -0.4163 0.02126 0.01375 -0.0388 0.9491 0.0654 -5.750 -0.3706 0.02041 0.01282 -0.0421 0.9435 0.0709 -5.500 -0.3281 0.01928 0.01175 -0.0448 0.9380 0.0767 -5.250 -0.2945 0.01870 0.01106 -0.0455 0.9283 0.0824 -5.000 -0.2611 0.01780 0.01027 -0.0463 0.9194 0.0890 -4.750 -0.2291 0.01714 0.00955 -0.0466 0.9094 0.0957 -4.500 -0.2038 0.01658 0.00906 -0.0458 0.8964 0.1023 -4.250 -0.1779 0.01600 0.00846 -0.0448 0.8833 0.1096 -4.000 -0.1529 0.01548 0.00798 -0.0438 0.8692 0.1176 -3.750 -0.1291 0.01490 0.00741 -0.0424 0.8541 0.1263 -3.500 -0.1051 0.01451 0.00697 -0.0411 0.8373 0.1369 -3.250 -0.0831 0.01396 0.00648 -0.0395 0.8184 0.1515 -3.000 -0.0615 0.01340 0.00600 -0.0379 0.7973 0.1724 -2.750 -0.0404 0.01281 0.00554 -0.0362 0.7747 0.2074 -2.500 -0.0224 0.01201 0.00508 -0.0342 0.7504 0.2864 -2.250 -0.0128 0.01090 0.00484 -0.0306 0.7247 0.5056 -2.000 0.0027 0.01045 0.00482 -0.0274 0.6967 0.6463 -1.750 0.0235 0.01035 0.00479 -0.0250 0.6639 0.7194 -1.500 0.0465 0.01039 0.00480 -0.0231 0.6250 0.7712 -1.250 0.0710 0.01064 0.00491 -0.0214 0.5822 0.8126 -1.000 0.0959 0.01101 0.00506 -0.0198 0.5376 0.8455 -0.750 0.1207 0.01143 0.00524 -0.0184 0.4972 0.8741 -0.500 0.1501 0.01195 0.00552 -0.0177 0.4661 0.8984 -0.250 0.1844 0.01252 0.00583 -0.0181 0.4435 0.9180 0.000 0.2277 0.01305 0.00611 -0.0205 0.4252 0.9301 0.250 0.2726 0.01350 0.00635 -0.0235 0.4104 0.9414 0.500 0.3260 0.01397 0.00663 -0.0281 0.3977 0.9516 0.750 0.3801 0.01441 0.00682 -0.0330 0.3873 0.9607 1.000 0.4193 0.01459 0.00691 -0.0354 0.3786 0.9696 1.250 0.4631 0.01484 0.00695 -0.0388 0.3707 0.9748 1.500 0.5022 0.01494 0.00701 -0.0412 0.3633 0.9823 1.750 0.5460 0.01505 0.00699 -0.0446 0.3568 0.9877 2.000 0.5873 0.01524 0.00708 -0.0476 0.3511 0.9939 2.250 0.6320 0.01525 0.00706 -0.0513 0.3448 0.9999 2.500 0.6520 0.01539 0.00712 -0.0502 0.3404 1.0000 2.750 0.6723 0.01570 0.00729 -0.0492 0.3362 1.0000 3.000 0.6917 0.01580 0.00744 -0.0478 0.3325 1.0000 3.250 0.7117 0.01596 0.00760 -0.0466 0.3285 1.0000 3.500 0.7322 0.01616 0.00776 -0.0454 0.3248 1.0000 3.750 0.7532 0.01642 0.00795 -0.0443 0.3215 1.0000 4.000 0.7746 0.01682 0.00828 -0.0433 0.3181 1.0000 4.250 0.7942 0.01701 0.00855 -0.0418 0.3146 1.0000 4.500 0.8143 0.01725 0.00880 -0.0405 0.3108 1.0000 4.750 0.8351 0.01752 0.00905 -0.0393 0.3075 1.0000 5.000 0.8565 0.01785 0.00932 -0.0382 0.3047 1.0000 5.250 0.8789 0.01841 0.00980 -0.0374 0.3017 1.0000 5.500 0.8978 0.01867 0.01017 -0.0358 0.2989 1.0000 5.750 0.9171 0.01896 0.01052 -0.0343 0.2954 1.0000 6.000 0.9371 0.01926 0.01085 -0.0330 0.2922 1.0000 6.250 0.9581 0.01960 0.01116 -0.0318 0.2894 1.0000 6.500 0.9801 0.02005 0.01155 -0.0308 0.2868 1.0000 6.750 1.0010 0.02065 0.01217 -0.0298 0.2842 1.0000 7.000 1.0188 0.02100 0.01264 -0.0280 0.2814 1.0000 7.250 1.0372 0.02137 0.01309 -0.0264 0.2782 1.0000 7.500 1.0570 0.02173 0.01348 -0.0251 0.2752 1.0000 7.750 1.0781 0.02211 0.01386 -0.0240 0.2726 1.0000 8.000 1.1010 0.02263 0.01432 -0.0232 0.2703 1.0000 8.250 1.1213 0.02334 0.01507 -0.0222 0.2677 1.0000 8.500 1.1377 0.02377 0.01567 -0.0203 0.2649 1.0000 8.750 1.1552 0.02422 0.01622 -0.0187 0.2618 1.0000 9.000 1.1745 0.02465 0.01670 -0.0174 0.2590 1.0000 9.250 1.1960 0.02508 0.01714 -0.0164 0.2566 1.0000 9.500 1.2200 0.02558 0.01759 -0.0160 0.2542 1.0000 9.750 1.2390 0.02636 0.01844 -0.0148 0.2515 1.0000 10.000 1.2531 0.02692 0.01919 -0.0129 0.2486 1.0000 10.250 1.2696 0.02753 0.01993 -0.0113 0.2457 1.0000 10.500 1.2881 0.02800 0.02047 -0.0100 0.2430 1.0000 10.750 1.3100 0.02835 0.02081 -0.0093 0.2404 1.0000 11.000 1.3362 0.02891 0.02129 -0.0093 0.2379 1.0000 11.250 1.3489 0.02980 0.02236 -0.0074 0.2353 1.0000 11.500 1.3590 0.03055 0.02332 -0.0051 0.2324 1.0000 11.750 1.3720 0.03115 0.02404 -0.0032 0.2294 1.0000 12.000 1.3892 0.03158 0.02452 -0.0019 0.2267 1.0000 12.250 1.4126 0.03189 0.02481 -0.0015 0.2243 1.0000 12.500 1.4377 0.03267 0.02554 -0.0015 0.2216 1.0000 12.750 1.4356 0.03362 0.02677 0.0022 0.2190 1.0000 13.000 1.4373 0.03453 0.02786 0.0053 0.2161 1.0000 13.250 1.4425 0.03518 0.02862 0.0082 0.2135 1.0000 13.500 1.4571 0.03548 0.02893 0.0097 0.2110 1.0000 13.750 1.4882 0.03563 0.02899 0.0091 0.2084 1.0000 14.000 1.4882 0.03681 0.03031 0.0121 0.2060 1.0000 14.250 1.4711 0.03837 0.03211 0.0166 0.2036 1.0000 14.500 1.4608 0.03998 0.03389 0.0196 0.2010 1.0000 14.750 1.4624 0.04100 0.03500 0.0215 0.1983 1.0000 15.000 1.4855 0.04097 0.03494 0.0219 0.1956 1.0000 15.250 1.5151 0.04108 0.03496 0.0217 0.1926 1.0000 15.500 1.4830 0.04415 0.03833 0.0250 0.1905 1.0000 15.750 1.4512 0.04794 0.04237 0.0269 0.1881 1.0000 16.000 1.4311 0.05134 0.04594 0.0275 0.1854 1.0000 16.250 1.4586 0.05066 0.04520 0.0279 0.1822 1.0000 16.500 1.4990 0.04921 0.04360 0.0282 0.1786 1.0000 16.750 1.4428 0.05629 0.05102 0.0280 0.1768 1.0000 17.000 0.8499 0.16022 0.15540 -0.0226 0.1422 1.0000 17.250 0.8790 0.15794 0.15316 -0.0209 0.1411 1.0000 17.500 1.4416 0.06291 0.05780 0.0267 0.1669 1.0000 17.750 1.3818 0.07318 0.06832 0.0228 0.1647 1.0000 18.000 0.6308 0.17718 0.17323 -0.0350 0.1419 1.0000