USA 51 AIRFOIL (usa51-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 51 AIRFOIL (usa51-il) Reynolds number: 200,000 Max Cl/Cd: 63.2 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa51-il-200000.txt Download as CSV file: xf-usa51-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: USA 51 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.4439 0.08620 0.08281 -0.0283 1.0000 0.0459
-8.250 -0.4581 0.08212 0.07882 -0.0325 1.0000 0.0466
-8.000 -0.4699 0.07806 0.07478 -0.0355 1.0000 0.0469
-7.750 -0.4800 0.07395 0.07057 -0.0378 1.0000 0.0472
-7.500 -0.4861 0.07043 0.06692 -0.0379 1.0000 0.0474
-7.250 -0.4930 0.06412 0.06068 -0.0371 1.0000 0.0484
-7.000 -0.4866 0.06182 0.05847 -0.0349 1.0000 0.0496
-6.750 -0.4819 0.05954 0.05619 -0.0332 1.0000 0.0510
-6.500 -0.4782 0.05672 0.05332 -0.0321 1.0000 0.0532
-6.250 -0.4744 0.05329 0.04972 -0.0314 1.0000 0.0564
-6.000 -0.4772 0.04821 0.04416 -0.0306 1.0000 0.0610
-5.750 -0.4666 0.04582 0.04188 -0.0290 1.0000 0.0628
-5.500 -0.4560 0.04399 0.04002 -0.0272 1.0000 0.0666
-5.250 -0.4495 0.04069 0.03631 -0.0256 1.0000 0.0751
-5.000 -0.4369 0.03909 0.03481 -0.0239 1.0000 0.0792
-4.750 -0.4276 0.03652 0.03186 -0.0221 1.0000 0.0889
-4.500 -0.4100 0.02733 0.02152 -0.0185 1.0000 0.0505
-4.250 -0.3848 0.02356 0.01717 -0.0172 0.9988 0.0423
-4.000 -0.3437 0.02098 0.01408 -0.0194 0.9942 0.0427
-3.750 -0.3044 0.01936 0.01208 -0.0213 0.9887 0.0453
-3.500 -0.2634 0.01795 0.01038 -0.0235 0.9842 0.0468
-3.250 -0.2273 0.01631 0.00870 -0.0251 0.9784 0.0515
-3.000 -0.1871 0.01544 0.00780 -0.0275 0.9725 0.0569
-2.750 -0.1495 0.01449 0.00684 -0.0294 0.9658 0.0634
-2.500 -0.1102 0.01388 0.00623 -0.0315 0.9586 0.0732
-2.250 -0.0734 0.01310 0.00558 -0.0333 0.9508 0.0961
-2.000 -0.0412 0.01142 0.00503 -0.0347 0.9434 0.3413
-1.750 -0.0219 0.00990 0.00496 -0.0327 0.9328 0.7059
-1.500 0.0512 0.00957 0.00514 -0.0401 0.9359 0.9332
-1.250 0.1327 0.00967 0.00507 -0.0503 0.9382 0.9813
-1.000 0.2052 0.00952 0.00477 -0.0592 0.9344 1.0000
-0.750 0.2437 0.00926 0.00441 -0.0613 0.9168 1.0000
-0.500 0.2803 0.00901 0.00406 -0.0629 0.8950 1.0000
-0.250 0.3170 0.00877 0.00369 -0.0643 0.8668 1.0000
0.000 0.3516 0.00860 0.00331 -0.0652 0.8284 1.0000
0.250 0.3774 0.00862 0.00309 -0.0643 0.7871 1.0000
0.500 0.3991 0.00873 0.00297 -0.0628 0.7472 1.0000
0.750 0.4194 0.00889 0.00290 -0.0610 0.7076 1.0000
1.000 0.4391 0.00907 0.00287 -0.0592 0.6678 1.0000
1.250 0.4588 0.00928 0.00285 -0.0574 0.6271 1.0000
1.500 0.4789 0.00950 0.00286 -0.0558 0.5908 1.0000
1.750 0.4996 0.00974 0.00291 -0.0543 0.5616 1.0000
2.000 0.5208 0.00999 0.00299 -0.0529 0.5372 1.0000
2.250 0.5425 0.01023 0.00309 -0.0517 0.5161 1.0000
2.500 0.5646 0.01049 0.00322 -0.0505 0.4988 1.0000
2.750 0.5872 0.01074 0.00338 -0.0495 0.4838 1.0000
3.000 0.6100 0.01098 0.00355 -0.0485 0.4696 1.0000
3.250 0.6329 0.01121 0.00373 -0.0475 0.4560 1.0000
3.500 0.6558 0.01144 0.00391 -0.0465 0.4422 1.0000
3.750 0.6788 0.01165 0.00411 -0.0455 0.4293 1.0000
4.000 0.7020 0.01187 0.00435 -0.0446 0.4172 1.0000
4.250 0.7251 0.01208 0.00457 -0.0436 0.4046 1.0000
4.500 0.7479 0.01228 0.00478 -0.0426 0.3907 1.0000
4.750 0.7705 0.01246 0.00499 -0.0416 0.3744 1.0000
5.000 0.7923 0.01264 0.00519 -0.0403 0.3519 1.0000
5.250 0.8127 0.01286 0.00532 -0.0389 0.3084 1.0000
5.500 0.8298 0.01347 0.00556 -0.0371 0.2322 1.0000
5.750 0.8467 0.01429 0.00605 -0.0354 0.1860 1.0000
6.000 0.8665 0.01489 0.00657 -0.0341 0.1677 1.0000
6.250 0.8868 0.01543 0.00707 -0.0329 0.1532 1.0000
6.500 0.9074 0.01594 0.00755 -0.0318 0.1394 1.0000
6.750 0.9286 0.01639 0.00799 -0.0308 0.1248 1.0000
7.000 0.9504 0.01682 0.00839 -0.0298 0.1080 1.0000
7.250 0.9723 0.01728 0.00876 -0.0287 0.0790 1.0000
7.500 0.9856 0.01853 0.00966 -0.0265 0.0385 1.0000
7.750 1.0032 0.01937 0.01065 -0.0246 0.0353 1.0000
8.000 1.0193 0.02034 0.01180 -0.0227 0.0329 1.0000
8.250 1.0320 0.02157 0.01318 -0.0203 0.0303 1.0000
8.500 1.0406 0.02303 0.01482 -0.0174 0.0287 1.0000
8.750 1.0526 0.02416 0.01607 -0.0150 0.0282 1.0000
9.000 1.0635 0.02536 0.01738 -0.0125 0.0277 1.0000
9.250 1.0735 0.02669 0.01879 -0.0099 0.0273 1.0000
9.500 1.0831 0.02811 0.02029 -0.0073 0.0269 1.0000
9.750 1.0942 0.02958 0.02184 -0.0050 0.0266 1.0000
10.000 1.1061 0.03114 0.02350 -0.0029 0.0258 1.0000
10.250 1.1184 0.03279 0.02525 -0.0010 0.0246 1.0000
10.500 1.1334 0.03464 0.02725 0.0005 0.0244 1.0000
10.750 1.1476 0.03677 0.02957 0.0021 0.0245 1.0000
11.000 1.1594 0.03906 0.03209 0.0038 0.0250 1.0000
11.250 1.1676 0.04162 0.03491 0.0058 0.0256 1.0000
11.500 1.1709 0.04462 0.03820 0.0081 0.0263 1.0000
11.750 1.1705 0.04815 0.04200 0.0102 0.0273 1.0000
12.000 1.1924 0.05010 0.04412 0.0112 0.0303 1.0000
12.250 1.1701 0.05253 0.04710 0.0151 0.0330 1.0000
12.500 1.1521 0.05694 0.05188 0.0166 0.0353 1.0000
12.750 1.0608 0.05243 0.04755 0.0191 0.0316 1.0000
13.000 1.0345 0.05794 0.05335 0.0183 0.0322 1.0000
13.250 1.0085 0.06426 0.05992 0.0165 0.0327 1.0000
13.500 0.9830 0.07104 0.06691 0.0139 0.0329 1.0000
13.750 0.9579 0.07810 0.07416 0.0107 0.0329 1.0000
14.000 0.9330 0.08540 0.08161 0.0068 0.0328 1.0000
14.250 0.9071 0.09306 0.08942 0.0024 0.0326 1.0000
14.500 0.8807 0.10086 0.09737 -0.0025 0.0323 1.0000
14.750 0.8525 0.10883 0.10546 -0.0078 0.0320 1.0000
15.000 0.8201 0.11660 0.11335 -0.0135 0.0317 1.0000
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Polar data table (+)
Polar graphs
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