NASA/LANGLEY RC08-64C AIRFOIL (rc0864c-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC08-64C AIRFOIL (rc0864c-il) Reynolds number: 50,000 Max Cl/Cd: 25.45 at α=3.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc0864c-il-50000-n5.txt Download as CSV file: xf-rc0864c-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC08-64C AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.4356 0.11333 0.10630 -0.0120 1.0000 0.1180
-9.750 -0.4518 0.11080 0.10384 -0.0149 1.0000 0.1229
-9.500 -0.4702 0.10792 0.10105 -0.0182 1.0000 0.1238
-9.250 -0.4551 0.10283 0.09599 -0.0169 1.0000 0.1250
-9.000 -0.4500 0.09869 0.09185 -0.0170 1.0000 0.1259
-8.500 -0.4514 0.08568 0.07878 -0.0226 1.0000 0.0605
-8.000 -0.4727 0.07498 0.06813 -0.0291 1.0000 0.0517
-7.750 -0.5694 0.08122 0.07393 -0.0262 1.0000 0.0526
-7.500 -0.5649 0.07741 0.07012 -0.0262 1.0000 0.0508
-7.250 -0.5627 0.07327 0.06594 -0.0268 1.0000 0.0488
-7.000 -0.5684 0.06725 0.05937 -0.0290 1.0000 0.0431
-6.750 -0.5590 0.06369 0.05582 -0.0283 1.0000 0.0425
-6.500 -0.5510 0.06020 0.05225 -0.0275 1.0000 0.0417
-6.250 -0.5426 0.05673 0.04862 -0.0266 1.0000 0.0407
-6.000 -0.5332 0.05326 0.04494 -0.0255 1.0000 0.0396
-5.750 -0.5227 0.04982 0.04119 -0.0242 1.0000 0.0383
-5.500 -0.5107 0.04638 0.03738 -0.0227 1.0000 0.0371
-5.250 -0.4967 0.04301 0.03353 -0.0209 1.0000 0.0357
-5.000 -0.4796 0.04014 0.03000 -0.0188 1.0000 0.0344
-4.750 -0.4619 0.03766 0.02718 -0.0173 1.0000 0.0340
-4.500 -0.4430 0.03530 0.02451 -0.0158 1.0000 0.0338
-4.250 -0.4226 0.03312 0.02199 -0.0144 1.0000 0.0336
-4.000 -0.4009 0.03113 0.01969 -0.0130 1.0000 0.0335
-3.750 -0.3781 0.02934 0.01761 -0.0117 1.0000 0.0336
-3.500 -0.3544 0.02779 0.01578 -0.0105 1.0000 0.0341
-3.250 -0.3310 0.02630 0.01420 -0.0094 1.0000 0.0355
-3.000 -0.3074 0.02514 0.01296 -0.0084 1.0000 0.0376
-2.750 -0.2832 0.02413 0.01179 -0.0073 1.0000 0.0396
-2.500 -0.2593 0.02323 0.01070 -0.0063 1.0000 0.0413
-2.250 -0.2361 0.02246 0.00973 -0.0052 1.0000 0.0430
-2.000 -0.2146 0.02171 0.00891 -0.0040 1.0000 0.0461
-1.750 -0.1927 0.02112 0.00822 -0.0028 1.0000 0.0520
-1.500 -0.0686 0.01709 0.00776 -0.0186 1.0000 1.0000
-1.250 -0.0562 0.01705 0.00749 -0.0159 1.0000 1.0000
-1.000 -0.0440 0.01705 0.00728 -0.0133 1.0000 1.0000
-0.750 -0.0317 0.01708 0.00713 -0.0108 1.0000 1.0000
-0.500 -0.0192 0.01714 0.00703 -0.0083 1.0000 1.0000
-0.250 -0.0064 0.01723 0.00699 -0.0060 1.0000 1.0000
0.000 0.0071 0.01736 0.00699 -0.0038 1.0000 1.0000
0.250 0.0213 0.01752 0.00705 -0.0018 1.0000 1.0000
0.500 0.0363 0.01771 0.00717 0.0000 1.0000 1.0000
0.750 0.0520 0.01795 0.00734 0.0016 1.0000 1.0000
1.000 0.0742 0.01823 0.00759 0.0018 0.9976 1.0000
1.250 0.1167 0.01861 0.00797 -0.0019 0.9871 1.0000
1.500 0.1589 0.01897 0.00838 -0.0054 0.9759 1.0000
1.750 0.2004 0.01931 0.00882 -0.0087 0.9636 1.0000
2.000 0.2413 0.01961 0.00927 -0.0118 0.9502 1.0000
2.250 0.2827 0.01987 0.00975 -0.0149 0.9359 1.0000
2.500 0.3371 0.01983 0.01002 -0.0196 0.9120 1.0000
2.750 0.4064 0.01872 0.00932 -0.0241 0.8514 1.0000
3.000 0.4442 0.01810 0.00893 -0.0231 0.7841 1.0000
3.250 0.4659 0.01831 0.00761 -0.0169 0.4469 1.0000
3.500 0.4665 0.02171 0.00847 -0.0133 0.0743 1.0000
3.750 0.4870 0.02286 0.00949 -0.0120 0.0533 1.0000
4.000 0.5079 0.02386 0.01055 -0.0106 0.0475 1.0000
4.250 0.5289 0.02482 0.01170 -0.0093 0.0439 1.0000
4.500 0.5494 0.02590 0.01290 -0.0079 0.0413 1.0000
4.750 0.5696 0.02719 0.01425 -0.0065 0.0391 1.0000
5.000 0.5918 0.02832 0.01559 -0.0051 0.0369 1.0000
5.250 0.6149 0.02960 0.01707 -0.0040 0.0348 1.0000
5.500 0.6387 0.03108 0.01870 -0.0029 0.0334 1.0000
5.750 0.6632 0.03281 0.02061 -0.0020 0.0327 1.0000
6.000 0.6873 0.03477 0.02281 -0.0010 0.0324 1.0000
6.250 0.7101 0.03694 0.02527 0.0000 0.0321 1.0000
6.500 0.7313 0.03934 0.02800 0.0012 0.0321 1.0000
6.750 0.7503 0.04199 0.03108 0.0027 0.0323 1.0000
7.000 0.7665 0.04493 0.03450 0.0043 0.0326 1.0000
7.250 0.7797 0.04814 0.03820 0.0060 0.0330 1.0000
7.500 0.7896 0.05160 0.04211 0.0078 0.0335 1.0000
7.750 0.7964 0.05523 0.04616 0.0095 0.0340 1.0000
8.000 0.8001 0.05898 0.05028 0.0110 0.0345 1.0000
8.250 0.8002 0.06288 0.05450 0.0124 0.0349 1.0000
8.500 0.7967 0.06690 0.05879 0.0136 0.0354 1.0000
8.750 0.7897 0.07100 0.06311 0.0144 0.0358 1.0000
9.000 0.7793 0.07516 0.06743 0.0149 0.0362 1.0000
9.250 0.7651 0.07934 0.07171 0.0150 0.0366 1.0000
9.500 0.7512 0.08410 0.07654 0.0135 0.0371 1.0000
9.750 0.7394 0.08945 0.08193 0.0106 0.0376 1.0000
10.000 0.7310 0.09514 0.08762 0.0072 0.0381 1.0000
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Polar data table (+)
Polar graphs
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