NASA/LANGLEY RC08-64C AIRFOIL (rc0864c-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC08-64C AIRFOIL (rc0864c-il) Reynolds number: 200,000 Max Cl/Cd: 44.01 at α=2.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc0864c-il-200000.txt Download as CSV file: xf-rc0864c-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.5632 0.08745 0.08382 -0.0172 1.0000 0.0374
-8.000 -0.5717 0.08191 0.07829 -0.0235 1.0000 0.0378
-7.750 -0.5818 0.07572 0.07195 -0.0297 1.0000 0.0383
-7.250 -0.5719 0.06920 0.06548 -0.0282 1.0000 0.0400
-7.000 -0.5669 0.06590 0.06212 -0.0284 1.0000 0.0416
-6.750 -0.5756 0.06137 0.05690 -0.0294 1.0000 0.0447
-6.500 -0.5621 0.05728 0.05315 -0.0287 1.0000 0.0458
-6.250 -0.5529 0.05499 0.05088 -0.0273 1.0000 0.0477
-6.000 -0.5516 0.05127 0.04661 -0.0255 1.0000 0.0526
-5.750 -0.5087 0.03469 0.03076 -0.0240 1.0000 0.0555
-5.500 -0.5122 0.03181 0.02732 -0.0206 1.0000 0.0613
-5.250 -0.5216 0.04317 0.03835 -0.0204 1.0000 0.0630
-5.000 -0.5087 0.04139 0.03647 -0.0186 1.0000 0.0681
-4.750 -0.4989 0.03870 0.03360 -0.0165 1.0000 0.0743
-4.500 -0.4873 0.03684 0.03147 -0.0143 1.0000 0.0847
-4.250 -0.4738 0.03592 0.03025 -0.0120 1.0000 0.0967
-4.000 -0.4585 0.03288 0.02741 -0.0110 1.0000 0.1007
-3.750 -0.4434 0.03113 0.02544 -0.0091 1.0000 0.1131
-3.500 -0.4270 0.02965 0.02377 -0.0075 1.0000 0.1273
-3.250 -0.4083 0.02792 0.02204 -0.0062 1.0000 0.1357
-3.000 -0.3899 0.02624 0.02027 -0.0049 1.0000 0.1509
-2.750 -0.3353 0.01845 0.01040 -0.0010 1.0000 0.0353
-2.500 -0.3096 0.01732 0.00905 0.0001 1.0000 0.0321
-2.250 -0.2841 0.01622 0.00791 0.0008 1.0000 0.0305
-2.000 -0.2595 0.01544 0.00707 0.0016 1.0000 0.0293
-1.750 -0.2359 0.01485 0.00646 0.0024 1.0000 0.0288
-1.500 -0.2125 0.01430 0.00594 0.0030 1.0000 0.0289
-1.250 -0.1790 0.01378 0.00544 0.0016 0.9975 0.0297
-1.000 -0.1383 0.01342 0.00505 -0.0012 0.9932 0.0336
-0.750 -0.0976 0.01308 0.00476 -0.0040 0.9883 0.0431
-0.500 -0.0752 0.01001 0.00480 -0.0034 0.9872 0.7909
-0.250 -0.0024 0.01026 0.00530 -0.0107 0.9940 1.0000
0.000 0.0429 0.01034 0.00526 -0.0146 0.9882 1.0000
0.250 0.0872 0.01040 0.00524 -0.0184 0.9818 1.0000
0.500 0.1343 0.01043 0.00524 -0.0227 0.9765 1.0000
0.750 0.1779 0.01042 0.00523 -0.0261 0.9686 1.0000
1.000 0.2349 0.01029 0.00514 -0.0322 0.9635 1.0000
1.250 0.3048 0.00949 0.00440 -0.0391 0.9362 1.0000
1.500 0.3367 0.00907 0.00399 -0.0386 0.9027 1.0000
1.750 0.3574 0.00878 0.00360 -0.0356 0.8480 1.0000
2.000 0.3782 0.00873 0.00334 -0.0330 0.7793 1.0000
2.250 0.3974 0.00903 0.00314 -0.0302 0.6619 1.0000
2.500 0.4085 0.01060 0.00320 -0.0269 0.3506 1.0000
2.750 0.4214 0.01277 0.00385 -0.0251 0.0438 1.0000
3.000 0.4434 0.01327 0.00431 -0.0239 0.0346 1.0000
3.500 0.4870 0.01421 0.00533 -0.0214 0.0286 1.0000
3.750 0.5079 0.01480 0.00596 -0.0200 0.0277 1.0000
4.000 0.5280 0.01550 0.00668 -0.0184 0.0272 1.0000
4.250 0.5478 0.01633 0.00749 -0.0168 0.0271 1.0000
4.500 0.5682 0.01727 0.00846 -0.0152 0.0273 1.0000
4.750 0.5897 0.01837 0.00959 -0.0137 0.0279 1.0000
5.000 0.6126 0.01969 0.01094 -0.0125 0.0288 1.0000
5.250 0.6369 0.02143 0.01264 -0.0116 0.0297 1.0000
5.500 0.6619 0.02292 0.01457 -0.0099 0.0329 1.0000
7.250 0.7380 0.03263 0.02794 0.0102 0.0793 1.0000
7.500 0.7934 0.05032 0.04546 0.0082 0.0703 1.0000
7.750 0.8076 0.05297 0.04809 0.0091 0.0674 1.0000
8.000 0.8220 0.05994 0.05477 0.0085 0.0653 1.0000
8.250 0.8084 0.06013 0.05582 0.0122 0.0594 1.0000
8.500 0.8209 0.06285 0.05849 0.0130 0.0574 1.0000
8.750 0.8329 0.06914 0.06458 0.0126 0.0559 1.0000
9.000 0.7944 0.07220 0.06833 0.0143 0.0541 1.0000
9.250 0.7747 0.07670 0.07296 0.0141 0.0532 1.0000
9.500 0.7551 0.08223 0.07857 0.0112 0.0528 1.0000
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Polar data table (+)
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