USA 51 AIRFOIL (usa51-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 51 AIRFOIL (usa51-il) Reynolds number: 500,000 Max Cl/Cd: 80.75 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa51-il-500000.txt Download as CSV file: xf-usa51-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.4528 0.11366 0.11130 -0.0170 1.0000 0.0183
-10.500 -0.4501 0.10996 0.10761 -0.0185 1.0000 0.0184
-10.250 -0.4484 0.10597 0.10364 -0.0204 1.0000 0.0186
-9.750 -0.3680 0.08563 0.08351 -0.0286 1.0000 0.0194
-9.500 -0.3647 0.08296 0.08085 -0.0281 1.0000 0.0199
-9.250 -0.3641 0.07966 0.07757 -0.0284 1.0000 0.0203
-9.000 -0.3641 0.07642 0.07435 -0.0287 1.0000 0.0208
-8.750 -0.3658 0.07291 0.07085 -0.0292 1.0000 0.0215
-8.500 -0.3719 0.06864 0.06661 -0.0301 1.0000 0.0217
-8.250 -0.3798 0.06445 0.06245 -0.0311 1.0000 0.0220
-8.000 -0.3922 0.06014 0.05819 -0.0321 1.0000 0.0219
-7.750 -0.4168 0.05602 0.05412 -0.0331 1.0000 0.0216
-7.500 -0.4372 0.05240 0.05050 -0.0329 1.0000 0.0217
-7.250 -0.4534 0.04765 0.04571 -0.0329 1.0000 0.0215
-7.000 -0.4624 0.04366 0.04166 -0.0322 1.0000 0.0219
-6.750 -0.4692 0.03950 0.03740 -0.0311 1.0000 0.0222
-6.500 -0.4631 0.03426 0.03199 -0.0323 0.9986 0.0233
-6.250 -0.4344 0.02910 0.02637 -0.0355 0.9953 0.0258
-6.000 -0.4778 0.02963 0.02586 -0.0311 0.9949 0.0182
-5.750 -0.4545 0.02369 0.01927 -0.0319 0.9908 0.0173
-5.500 -0.4235 0.02044 0.01551 -0.0328 0.9871 0.0176
-5.250 -0.3873 0.01846 0.01314 -0.0345 0.9844 0.0182
-5.000 -0.3549 0.01629 0.01065 -0.0357 0.9810 0.0199
-4.750 -0.3212 0.01526 0.00955 -0.0370 0.9762 0.0211
-4.500 -0.2833 0.01427 0.00846 -0.0391 0.9728 0.0226
-4.250 -0.2448 0.01341 0.00750 -0.0412 0.9706 0.0245
-4.000 -0.2119 0.01312 0.00713 -0.0420 0.9653 0.0259
-3.750 -0.1806 0.01151 0.00546 -0.0428 0.9607 0.0295
-3.500 -0.1423 0.01087 0.00480 -0.0449 0.9576 0.0328
-3.250 -0.1091 0.01046 0.00434 -0.0459 0.9516 0.0355
-3.000 -0.0781 0.00970 0.00354 -0.0464 0.9441 0.0403
-2.750 -0.0486 0.00930 0.00311 -0.0466 0.9339 0.0446
-2.500 -0.0182 0.00894 0.00272 -0.0469 0.9225 0.0509
-2.250 0.0108 0.00852 0.00239 -0.0469 0.9068 0.0747
-2.000 0.0387 0.00783 0.00211 -0.0470 0.8866 0.1937
-1.750 0.0657 0.00725 0.00188 -0.0469 0.8610 0.3144
-1.500 0.0865 0.00646 0.00171 -0.0456 0.8331 0.5281
-1.250 0.1043 0.00596 0.00166 -0.0432 0.8047 0.7053
-1.000 0.1246 0.00577 0.00166 -0.0411 0.7762 0.8001
-0.750 0.1485 0.00573 0.00168 -0.0397 0.7474 0.8709
-0.500 0.1848 0.00581 0.00171 -0.0410 0.7182 0.9208
-0.250 0.2272 0.00599 0.00174 -0.0439 0.6885 0.9457
0.000 0.2700 0.00623 0.00180 -0.0468 0.6521 0.9654
0.250 0.3145 0.00656 0.00187 -0.0502 0.6000 0.9805
0.500 0.3589 0.00693 0.00190 -0.0538 0.5349 0.9894
0.750 0.3996 0.00721 0.00193 -0.0567 0.4952 0.9956
1.000 0.4404 0.00737 0.00193 -0.0597 0.4706 1.0000
1.250 0.4614 0.00748 0.00194 -0.0585 0.4553 1.0000
1.500 0.4830 0.00758 0.00196 -0.0573 0.4419 1.0000
1.750 0.5050 0.00768 0.00200 -0.0561 0.4295 1.0000
2.000 0.5274 0.00779 0.00205 -0.0551 0.4195 1.0000
2.250 0.5499 0.00792 0.00211 -0.0540 0.4096 1.0000
2.500 0.5730 0.00802 0.00218 -0.0531 0.3991 1.0000
2.750 0.5963 0.00813 0.00227 -0.0521 0.3886 1.0000
3.000 0.6195 0.00826 0.00235 -0.0512 0.3772 1.0000
3.250 0.6428 0.00839 0.00244 -0.0503 0.3640 1.0000
3.500 0.6663 0.00852 0.00254 -0.0494 0.3510 1.0000
3.750 0.6898 0.00866 0.00265 -0.0485 0.3356 1.0000
4.000 0.7130 0.00883 0.00277 -0.0476 0.3145 1.0000
4.250 0.7341 0.00920 0.00292 -0.0464 0.2660 1.0000
4.500 0.7489 0.01019 0.00333 -0.0443 0.1619 1.0000
4.750 0.7697 0.01065 0.00366 -0.0430 0.1398 1.0000
5.000 0.7917 0.01100 0.00398 -0.0419 0.1303 1.0000
5.250 0.8145 0.01127 0.00424 -0.0410 0.1236 1.0000
5.500 0.8368 0.01160 0.00454 -0.0400 0.1167 1.0000
5.750 0.8601 0.01182 0.00478 -0.0392 0.1100 1.0000
6.000 0.8830 0.01208 0.00504 -0.0383 0.1005 1.0000
6.250 0.9021 0.01271 0.00531 -0.0369 0.0456 1.0000
6.500 0.9200 0.01348 0.00604 -0.0351 0.0270 1.0000
6.750 0.9405 0.01401 0.00668 -0.0337 0.0239 1.0000
7.000 0.9593 0.01471 0.00753 -0.0320 0.0209 1.0000
7.250 0.9789 0.01532 0.00823 -0.0305 0.0194 1.0000
7.500 0.9982 0.01593 0.00891 -0.0290 0.0183 1.0000
7.750 1.0163 0.01663 0.00970 -0.0274 0.0173 1.0000
8.000 1.0332 0.01742 0.01057 -0.0255 0.0164 1.0000
8.250 1.0484 0.01832 0.01153 -0.0235 0.0156 1.0000
8.500 1.0597 0.01949 0.01279 -0.0209 0.0148 1.0000
9.000 1.0770 0.02231 0.01577 -0.0152 0.0136 1.0000
9.250 1.0923 0.02310 0.01664 -0.0133 0.0131 1.0000
9.500 1.1047 0.02419 0.01781 -0.0110 0.0127 1.0000
9.750 1.1154 0.02534 0.01903 -0.0085 0.0124 1.0000
10.000 1.1255 0.02662 0.02039 -0.0060 0.0121 1.0000
10.250 1.1360 0.02798 0.02185 -0.0036 0.0118 1.0000
10.500 1.1473 0.02955 0.02353 -0.0015 0.0117 1.0000
10.750 1.1588 0.03136 0.02545 0.0004 0.0116 1.0000
11.000 1.1694 0.03326 0.02748 0.0023 0.0115 1.0000
11.250 1.1780 0.03531 0.02970 0.0044 0.0115 1.0000
11.500 1.1845 0.03657 0.03099 0.0063 0.0108 1.0000
11.750 1.1916 0.03825 0.03274 0.0080 0.0105 1.0000
12.000 1.1960 0.04037 0.03494 0.0097 0.0102 1.0000
12.250 1.1991 0.04287 0.03757 0.0114 0.0101 1.0000
17.000 0.6977 0.17533 0.17335 -0.0511 0.0201 1.0000
17.250 0.6968 0.18060 0.17862 -0.0538 0.0200 1.0000
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Polar data table (+)
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