NASA/LANGLEY RC10-64C AIRFOIL (rc1064c-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC10-64C AIRFOIL (rc1064c-il) Reynolds number: 200,000 Max Cl/Cd: 57.69 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc1064c-il-200000.txt Download as CSV file: xf-rc1064c-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC10-64C 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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.4454 0.10328 0.09969 -0.0239 1.0000 0.0636
-10.500 -0.4479 0.09922 0.09565 -0.0249 1.0000 0.0655
-10.250 -0.5876 0.09770 0.09402 -0.0245 1.0000 0.0599
-10.000 -0.5759 0.09605 0.09236 -0.0224 1.0000 0.0618
-9.750 -0.5748 0.09241 0.08875 -0.0236 1.0000 0.0636
-9.500 -0.5797 0.08764 0.08400 -0.0265 1.0000 0.0653
-9.250 -0.5931 0.08140 0.07781 -0.0321 1.0000 0.0664
-9.000 -0.6138 0.07620 0.07258 -0.0348 1.0000 0.0672
-8.750 -0.6364 0.07181 0.06812 -0.0346 1.0000 0.0686
-8.500 -0.6729 0.06924 0.06510 -0.0320 1.0000 0.0708
-8.000 -0.6778 0.05909 0.05492 -0.0288 1.0000 0.0731
-7.750 -0.6706 0.05648 0.05232 -0.0269 1.0000 0.0743
-7.500 -0.6663 0.05399 0.04977 -0.0246 1.0000 0.0759
-7.250 -0.7056 0.03953 0.03365 -0.0149 1.0000 0.0472
-7.000 -0.6955 0.03572 0.02973 -0.0126 1.0000 0.0453
-6.750 -0.6873 0.03236 0.02602 -0.0097 1.0000 0.0445
-6.500 -0.6751 0.02994 0.02322 -0.0070 1.0000 0.0451
-6.250 -0.6602 0.02780 0.02072 -0.0047 1.0000 0.0456
-6.000 -0.6434 0.02566 0.01824 -0.0026 1.0000 0.0457
-5.750 -0.6246 0.02396 0.01624 -0.0009 1.0000 0.0461
-5.500 -0.6047 0.02271 0.01472 0.0007 1.0000 0.0467
-5.250 -0.5832 0.02058 0.01242 0.0018 1.0000 0.0481
-5.000 -0.5626 0.01965 0.01150 0.0029 1.0000 0.0507
-4.750 -0.5414 0.01894 0.01073 0.0041 1.0000 0.0535
-4.500 -0.5193 0.01816 0.00985 0.0052 1.0000 0.0560
-4.250 -0.4973 0.01737 0.00898 0.0063 1.0000 0.0588
-4.000 -0.4766 0.01649 0.00818 0.0073 1.0000 0.0641
-3.750 -0.4553 0.01613 0.00776 0.0084 1.0000 0.0714
-3.500 -0.4347 0.01544 0.00716 0.0095 1.0000 0.0832
-3.250 -0.4073 0.01473 0.00658 0.0091 0.9986 0.1053
-3.000 -0.3704 0.01417 0.00630 0.0067 0.9952 0.1501
-2.750 -0.3339 0.01384 0.00604 0.0045 0.9910 0.2062
-2.500 -0.3016 0.01264 0.00567 0.0025 0.9873 0.3461
-2.250 -0.2785 0.01101 0.00580 0.0035 0.9844 0.7348
-2.000 -0.2490 0.01105 0.00617 0.0044 0.9792 0.8523
-1.750 -0.2104 0.01136 0.00646 0.0031 0.9757 0.9036
-1.500 -0.1696 0.01160 0.00664 0.0009 0.9705 0.9290
-1.250 -0.1193 0.01190 0.00684 -0.0032 0.9676 0.9490
-1.000 -0.0561 0.01234 0.00720 -0.0098 0.9680 0.9685
-0.750 0.0144 0.01260 0.00738 -0.0184 0.9697 0.9803
-0.500 0.0730 0.01258 0.00732 -0.0250 0.9682 0.9864
-0.250 0.1273 0.01247 0.00718 -0.0308 0.9660 0.9905
0.000 0.1781 0.01233 0.00702 -0.0359 0.9624 0.9938
0.250 0.2320 0.01207 0.00676 -0.0415 0.9586 0.9953
0.500 0.2869 0.01175 0.00646 -0.0472 0.9550 0.9966
0.750 0.3385 0.01145 0.00619 -0.0522 0.9481 0.9982
1.000 0.3880 0.01122 0.00598 -0.0564 0.9373 0.9997
1.250 0.4137 0.01138 0.00613 -0.0556 0.9166 1.0000
1.500 0.4236 0.01126 0.00607 -0.0520 0.8878 1.0000
1.750 0.4443 0.01089 0.00572 -0.0501 0.8692 1.0000
2.000 0.4669 0.01049 0.00529 -0.0483 0.8526 1.0000
2.250 0.4911 0.01027 0.00501 -0.0469 0.8337 1.0000
2.500 0.5125 0.01022 0.00492 -0.0452 0.8044 1.0000
2.750 0.5334 0.01016 0.00480 -0.0434 0.7692 1.0000
3.000 0.5548 0.01016 0.00471 -0.0417 0.7327 1.0000
3.250 0.5763 0.01022 0.00467 -0.0401 0.6930 1.0000
3.500 0.5968 0.01037 0.00465 -0.0384 0.6420 1.0000
3.750 0.6155 0.01067 0.00468 -0.0363 0.5718 1.0000
4.000 0.6312 0.01121 0.00478 -0.0338 0.4741 1.0000
4.250 0.6427 0.01215 0.00506 -0.0309 0.3400 1.0000
4.500 0.6528 0.01341 0.00557 -0.0281 0.2031 1.0000
4.750 0.6652 0.01454 0.00621 -0.0256 0.1305 1.0000
5.000 0.6801 0.01538 0.00688 -0.0233 0.1046 1.0000
5.250 0.6968 0.01603 0.00748 -0.0212 0.0926 1.0000
5.500 0.7132 0.01672 0.00817 -0.0191 0.0847 1.0000
5.750 0.7299 0.01740 0.00880 -0.0171 0.0783 1.0000
6.000 0.7463 0.01832 0.00972 -0.0150 0.0741 1.0000
6.250 0.7651 0.01901 0.01049 -0.0132 0.0705 1.0000
6.500 0.7838 0.01974 0.01121 -0.0116 0.0669 1.0000
6.750 0.8030 0.02133 0.01273 -0.0103 0.0629 1.0000
7.000 0.8236 0.02205 0.01358 -0.0088 0.0611 1.0000
7.250 0.8447 0.02307 0.01474 -0.0075 0.0591 1.0000
7.500 0.8656 0.02415 0.01593 -0.0062 0.0569 1.0000
7.750 0.8855 0.02514 0.01698 -0.0049 0.0545 1.0000
8.000 0.9059 0.02673 0.01860 -0.0040 0.0523 1.0000
8.250 0.9251 0.02978 0.02190 -0.0028 0.0509 1.0000
8.500 0.9415 0.03112 0.02355 -0.0007 0.0500 1.0000
8.750 0.9564 0.03268 0.02543 0.0015 0.0486 1.0000
9.000 0.9698 0.03454 0.02761 0.0037 0.0470 1.0000
9.250 0.9803 0.03709 0.03050 0.0062 0.0462 1.0000
9.500 0.9872 0.04010 0.03389 0.0089 0.0458 1.0000
9.750 0.9911 0.04313 0.03726 0.0118 0.0452 1.0000
10.000 0.9977 0.04536 0.03971 0.0141 0.0440 1.0000
10.250 1.0159 0.04639 0.04059 0.0145 0.0419 1.0000
10.500 1.0134 0.05014 0.04462 0.0172 0.0415 1.0000
10.750 0.9395 0.05991 0.05546 0.0263 0.0466 1.0000
11.000 0.9137 0.06440 0.06015 0.0291 0.0478 1.0000
11.250 0.8902 0.06911 0.06500 0.0302 0.0488 1.0000
11.500 0.8712 0.07444 0.07040 0.0298 0.0500 1.0000
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Polar data table (+)
Polar graphs
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