XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY RC-12(B)3 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -10.250 -0.5016 0.09202 0.08686 -0.0362 1.0000 0.0388 -10.000 -0.5088 0.08621 0.08109 -0.0400 1.0000 0.0386 -9.750 -0.5219 0.07935 0.07427 -0.0456 1.0000 0.0382 -9.500 -0.5407 0.07321 0.06811 -0.0485 1.0000 0.0379 -9.250 -0.5633 0.06834 0.06318 -0.0480 1.0000 0.0377 -9.000 -0.5872 0.06435 0.05913 -0.0451 1.0000 0.0375 -8.750 -0.6123 0.06111 0.05581 -0.0403 1.0000 0.0373 -8.500 -0.6368 0.05817 0.05274 -0.0346 1.0000 0.0373 -8.250 -0.6545 0.05455 0.04891 -0.0301 0.9984 0.0373 -8.000 -0.6530 0.04849 0.04232 -0.0297 0.9903 0.0377 -7.750 -0.6478 0.04269 0.03577 -0.0287 0.9834 0.0391 -7.500 -0.6376 0.03856 0.03095 -0.0272 0.9761 0.0403 -7.250 -0.6123 0.03666 0.02889 -0.0280 0.9720 0.0414 -7.000 -0.5909 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-0.0533 0.7157 1.0000 2.750 0.6185 0.01640 0.00804 -0.0523 0.7024 1.0000 3.000 0.6420 0.01640 0.00803 -0.0513 0.6885 1.0000 3.250 0.6654 0.01640 0.00804 -0.0503 0.6737 1.0000 3.500 0.6887 0.01642 0.00806 -0.0493 0.6574 1.0000 3.750 0.7118 0.01644 0.00808 -0.0482 0.6394 1.0000 4.000 0.7349 0.01648 0.00809 -0.0471 0.6201 1.0000 4.250 0.7575 0.01656 0.00815 -0.0460 0.5992 1.0000 4.500 0.7800 0.01668 0.00827 -0.0449 0.5772 1.0000 4.750 0.8023 0.01683 0.00837 -0.0437 0.5547 1.0000 5.000 0.8240 0.01704 0.00853 -0.0425 0.5295 1.0000 5.250 0.8450 0.01729 0.00871 -0.0411 0.5023 1.0000 5.500 0.8653 0.01761 0.00895 -0.0397 0.4741 1.0000 5.750 0.8848 0.01798 0.00923 -0.0382 0.4458 1.0000 6.000 0.9037 0.01840 0.00956 -0.0366 0.4180 1.0000 6.250 0.9219 0.01886 0.00995 -0.0350 0.3904 1.0000 6.500 0.9393 0.01936 0.01039 -0.0333 0.3636 1.0000 6.750 0.9556 0.01992 0.01086 -0.0314 0.3372 1.0000 7.000 0.9712 0.02050 0.01137 -0.0294 0.3111 1.0000 7.250 0.9860 0.02111 0.01191 -0.0273 0.2855 1.0000 7.500 0.9994 0.02178 0.01251 -0.0251 0.2615 1.0000 7.750 1.0124 0.02245 0.01314 -0.0228 0.2379 1.0000 8.000 1.0239 0.02318 0.01380 -0.0202 0.2161 1.0000 8.250 1.0349 0.02392 0.01449 -0.0177 0.1952 1.0000 8.500 1.0442 0.02472 0.01523 -0.0148 0.1763 1.0000 8.750 1.0523 0.02556 0.01603 -0.0119 0.1599 1.0000 9.000 1.0592 0.02642 0.01685 -0.0087 0.1452 1.0000 9.250 1.0654 0.02729 0.01771 -0.0055 0.1321 1.0000 9.500 1.0692 0.02812 0.01856 -0.0018 0.1211 1.0000 9.750 1.0704 0.02901 0.01946 0.0022 0.1120 1.0000 10.000 1.0702 0.03002 0.02044 0.0061 0.1042 1.0000 10.250 1.0737 0.03100 0.02151 0.0094 0.0964 1.0000 10.500 1.0735 0.03223 0.02268 0.0128 0.0906 1.0000 10.750 1.0783 0.03333 0.02390 0.0156 0.0841 1.0000 11.000 1.0793 0.03467 0.02524 0.0185 0.0792 1.0000 11.250 1.0829 0.03599 0.02665 0.0210 0.0746 1.0000 11.500 1.0863 0.03736 0.02810 0.0234 0.0702 1.0000 11.750 1.0867 0.03899 0.02971 0.0258 0.0669 1.0000 12.000 1.0912 0.04049 0.03136 0.0278 0.0635 1.0000 12.250 1.0949 0.04209 0.03307 0.0297 0.0604 1.0000 12.500 1.0970 0.04382 0.03486 0.0315 0.0578 1.0000 12.750 1.0981 0.04575 0.03679 0.0331 0.0555 1.0000 13.000 1.1023 0.04757 0.03881 0.0346 0.0530 1.0000 13.250 1.1050 0.04954 0.04092 0.0359 0.0507 1.0000 13.500 1.1068 0.05161 0.04308 0.0370 0.0488 1.0000 13.750 1.1077 0.05382 0.04532 0.0380 0.0473 1.0000 14.000 1.1089 0.05617 0.04779 0.0390 0.0457 1.0000 14.250 1.1083 0.05880 0.05066 0.0397 0.0440 1.0000 14.500 1.1066 0.06157 0.05361 0.0402 0.0424 1.0000 14.750 1.1041 0.06447 0.05666 0.0405 0.0411 1.0000 15.000 1.1015 0.06746 0.05975 0.0406 0.0399 1.0000 15.250 1.0991 0.07048 0.06284 0.0405 0.0389 1.0000 15.500 1.0969 0.07361 0.06601 0.0404 0.0379 1.0000 15.750 1.0861 0.07814 0.07082 0.0394 0.0371 1.0000 16.000 1.0736 0.08314 0.07608 0.0380 0.0363 1.0000 16.250 1.0596 0.08860 0.08178 0.0361 0.0357 1.0000 16.500 1.0438 0.09462 0.08803 0.0336 0.0352 1.0000 16.750 1.0254 0.10144 0.09506 0.0304 0.0348 1.0000 17.000 1.0031 0.10940 0.10323 0.0264 0.0346 1.0000 17.250 0.9729 0.11960 0.11366 0.0208 0.0347 1.0000 17.500 0.9196 0.13640 0.13070 0.0111 0.0358 1.0000