USA 51 AIRFOIL (usa51-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: USA 51 AIRFOIL (usa51-il) Reynolds number: 200,000 Max Cl/Cd: 60.29 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa51-il-200000-n5.txt Download as CSV file: xf-usa51-il-200000-n5.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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.4386 0.08487 0.08145 -0.0277 1.0000 0.0134
-8.500 -0.4446 0.08072 0.07735 -0.0292 1.0000 0.0132
-8.250 -0.4528 0.07681 0.07351 -0.0305 1.0000 0.0129
-8.000 -0.4671 0.07294 0.06975 -0.0317 1.0000 0.0128
-7.750 -0.4773 0.06809 0.06492 -0.0340 1.0000 0.0125
-7.500 -0.4858 0.06296 0.05976 -0.0355 1.0000 0.0124
-7.250 -0.4924 0.05773 0.05446 -0.0361 1.0000 0.0123
-7.000 -0.4968 0.05233 0.04891 -0.0359 1.0000 0.0122
-6.750 -0.4999 0.04641 0.04276 -0.0350 1.0000 0.0120
-6.500 -0.5007 0.04039 0.03641 -0.0334 1.0000 0.0119
-6.250 -0.4948 0.03634 0.03203 -0.0315 1.0000 0.0124
-6.000 -0.4686 0.03170 0.02685 -0.0332 0.9958 0.0137
-5.750 -0.4402 0.02715 0.02157 -0.0346 0.9905 0.0143
-5.500 -0.4081 0.02379 0.01757 -0.0360 0.9861 0.0149
-5.250 -0.3772 0.02162 0.01491 -0.0367 0.9805 0.0156
-5.000 -0.3435 0.01941 0.01239 -0.0384 0.9765 0.0172
-4.750 -0.3122 0.01866 0.01154 -0.0393 0.9697 0.0197
-4.500 -0.2759 0.01732 0.00999 -0.0409 0.9653 0.0216
-4.250 -0.2451 0.01637 0.00884 -0.0413 0.9572 0.0242
-4.000 -0.2085 0.01523 0.00761 -0.0432 0.9519 0.0281
-3.750 -0.1798 0.01447 0.00679 -0.0433 0.9415 0.0317
-3.500 -0.1483 0.01396 0.00617 -0.0439 0.9329 0.0366
-3.250 -0.1175 0.01319 0.00536 -0.0445 0.9242 0.0420
-3.000 -0.0892 0.01278 0.00488 -0.0445 0.9117 0.0478
-2.750 -0.0594 0.01231 0.00439 -0.0447 0.8988 0.0544
-2.500 -0.0280 0.01186 0.00397 -0.0454 0.8849 0.0727
-2.250 0.0048 0.01133 0.00359 -0.0463 0.8694 0.1245
-2.000 0.0379 0.01087 0.00329 -0.0474 0.8517 0.1903
-1.750 0.0690 0.01031 0.00299 -0.0482 0.8311 0.2843
-1.500 0.0965 0.00971 0.00277 -0.0482 0.8100 0.4218
-1.250 0.1161 0.00894 0.00270 -0.0462 0.7884 0.6339
-1.000 0.1405 0.00853 0.00273 -0.0446 0.7672 0.8016
-0.750 0.1903 0.00851 0.00276 -0.0483 0.7434 0.9054
-0.500 0.2423 0.00864 0.00272 -0.0529 0.7181 0.9472
-0.250 0.2881 0.00882 0.00269 -0.0564 0.6880 0.9738
0.000 0.3306 0.00899 0.00261 -0.0594 0.6477 0.9893
0.250 0.3710 0.00920 0.00250 -0.0622 0.5962 1.0000
0.500 0.3903 0.00940 0.00244 -0.0605 0.5544 1.0000
0.750 0.4105 0.00960 0.00242 -0.0590 0.5224 1.0000
1.000 0.4315 0.00979 0.00244 -0.0577 0.4976 1.0000
1.500 0.4749 0.01018 0.00255 -0.0553 0.4613 1.0000
1.750 0.4973 0.01037 0.00262 -0.0543 0.4474 1.0000
2.000 0.5201 0.01055 0.00272 -0.0533 0.4354 1.0000
2.250 0.5430 0.01073 0.00284 -0.0524 0.4230 1.0000
2.500 0.5659 0.01092 0.00296 -0.0514 0.4098 1.0000
2.750 0.5891 0.01110 0.00309 -0.0505 0.3977 1.0000
3.000 0.6124 0.01129 0.00324 -0.0496 0.3874 1.0000
3.250 0.6357 0.01147 0.00340 -0.0487 0.3755 1.0000
3.500 0.6591 0.01165 0.00356 -0.0479 0.3627 1.0000
3.750 0.6824 0.01184 0.00373 -0.0470 0.3477 1.0000
4.000 0.7055 0.01205 0.00392 -0.0461 0.3315 1.0000
4.250 0.7288 0.01225 0.00412 -0.0452 0.3146 1.0000
4.500 0.7516 0.01249 0.00433 -0.0443 0.2925 1.0000
4.750 0.7735 0.01283 0.00457 -0.0432 0.2610 1.0000
5.000 0.7934 0.01336 0.00490 -0.0419 0.2115 1.0000
5.250 0.8127 0.01401 0.00534 -0.0406 0.1718 1.0000
5.500 0.8332 0.01457 0.00580 -0.0394 0.1527 1.0000
5.750 0.8535 0.01514 0.00629 -0.0383 0.1366 1.0000
6.000 0.8746 0.01562 0.00677 -0.0372 0.1250 1.0000
6.250 0.8960 0.01607 0.00724 -0.0362 0.1162 1.0000
6.500 0.9170 0.01656 0.00770 -0.0352 0.1039 1.0000
6.750 0.9392 0.01694 0.00808 -0.0343 0.0847 1.0000
7.000 0.9523 0.01821 0.00899 -0.0322 0.0282 1.0000
7.250 0.9697 0.01907 0.01000 -0.0304 0.0225 1.0000
7.500 0.9866 0.02000 0.01109 -0.0286 0.0194 1.0000
7.750 1.0054 0.02071 0.01197 -0.0272 0.0180 1.0000
8.000 1.0231 0.02150 0.01291 -0.0256 0.0160 1.0000
8.250 1.0392 0.02241 0.01396 -0.0239 0.0145 1.0000
8.500 1.0529 0.02346 0.01517 -0.0219 0.0136 1.0000
8.750 1.0626 0.02473 0.01659 -0.0193 0.0128 1.0000
9.000 1.0664 0.02631 0.01829 -0.0161 0.0122 1.0000
9.250 1.0657 0.02798 0.02003 -0.0123 0.0116 1.0000
9.500 1.0759 0.02891 0.02112 -0.0099 0.0110 1.0000
9.750 1.0826 0.03014 0.02244 -0.0074 0.0105 1.0000
10.000 1.0886 0.03152 0.02391 -0.0049 0.0100 1.0000
10.250 1.0939 0.03313 0.02560 -0.0026 0.0097 1.0000
10.500 1.0996 0.03484 0.02740 -0.0006 0.0094 1.0000
10.750 1.1058 0.03660 0.02926 0.0012 0.0092 1.0000
11.000 1.1121 0.03854 0.03130 0.0028 0.0089 1.0000
11.250 1.1181 0.04049 0.03336 0.0042 0.0087 1.0000
11.500 1.1235 0.04268 0.03567 0.0054 0.0085 1.0000
11.750 1.1271 0.04502 0.03814 0.0065 0.0084 1.0000
12.000 1.1292 0.04750 0.04074 0.0074 0.0082 1.0000
12.250 1.1306 0.05026 0.04365 0.0081 0.0081 1.0000
12.500 1.1271 0.05378 0.04730 0.0085 0.0076 1.0000
12.750 1.1239 0.05696 0.05067 0.0086 0.0076 1.0000
13.000 1.1174 0.06048 0.05441 0.0081 0.0075 1.0000
13.250 1.1109 0.06443 0.05855 0.0073 0.0075 1.0000
13.500 1.0999 0.06885 0.06323 0.0056 0.0073 1.0000
13.750 1.0882 0.07377 0.06838 0.0033 0.0070 1.0000
14.000 1.0784 0.07876 0.07355 0.0010 0.0070 1.0000
14.250 1.0665 0.08435 0.07933 -0.0018 0.0070 1.0000
14.500 1.0547 0.09015 0.08528 -0.0048 0.0071 1.0000
14.750 1.0396 0.09688 0.09221 -0.0087 0.0070 1.0000
15.000 1.0285 0.10303 0.09849 -0.0122 0.0071 1.0000
15.250 1.0148 0.11004 0.10565 -0.0163 0.0071 1.0000
15.500 0.9985 0.11807 0.11385 -0.0212 0.0071 1.0000
15.750 0.9861 0.12549 0.12139 -0.0257 0.0072 1.0000
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Polar data table (+)
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