NASA/LANGLEY RC10-64C AIRFOIL (rc1064c-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC10-64C AIRFOIL (rc1064c-il) Reynolds number: 50,000 Max Cl/Cd: 30.6 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc1064c-il-50000-n5.txt Download as CSV file: xf-rc1064c-il-50000-n5.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 = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5860 0.09582 0.08832 -0.0279 1.0000 0.0600
-9.750 -0.5927 0.09007 0.08260 -0.0316 1.0000 0.0595
-9.500 -0.6033 0.08462 0.07721 -0.0346 1.0000 0.0590
-9.250 -0.6173 0.07986 0.07245 -0.0358 1.0000 0.0586
-9.000 -0.6317 0.07552 0.06807 -0.0357 1.0000 0.0581
-8.750 -0.6435 0.07115 0.06361 -0.0352 1.0000 0.0577
-8.500 -0.6531 0.06688 0.05919 -0.0342 1.0000 0.0574
-8.250 -0.6602 0.06274 0.05485 -0.0326 1.0000 0.0571
-8.000 -0.6646 0.05878 0.05064 -0.0307 1.0000 0.0569
-7.500 -0.6642 0.05148 0.04268 -0.0260 1.0000 0.0572
-7.250 -0.6598 0.04809 0.03892 -0.0236 1.0000 0.0576
-7.000 -0.6527 0.04493 0.03533 -0.0211 1.0000 0.0583
-6.750 -0.6431 0.04205 0.03195 -0.0187 1.0000 0.0601
-6.500 -0.6310 0.03945 0.02876 -0.0163 1.0000 0.0623
-6.250 -0.6156 0.03737 0.02661 -0.0148 1.0000 0.0646
-6.000 -0.5982 0.03541 0.02441 -0.0132 1.0000 0.0666
-5.750 -0.5791 0.03346 0.02217 -0.0117 1.0000 0.0689
-5.500 -0.5581 0.03162 0.01998 -0.0103 1.0000 0.0718
-5.250 -0.5360 0.03004 0.01808 -0.0090 1.0000 0.0766
-5.000 -0.5154 0.02886 0.01694 -0.0079 1.0000 0.0827
-4.750 -0.4910 0.02758 0.01538 -0.0069 1.0000 0.0897
-4.250 -0.4460 0.02551 0.01334 -0.0049 1.0000 0.1117
-4.000 -0.4241 0.02466 0.01250 -0.0039 1.0000 0.1291
-3.750 -0.4030 0.02393 0.01181 -0.0027 1.0000 0.1511
-3.500 -0.3821 0.02335 0.01114 -0.0015 1.0000 0.1835
-3.250 -0.3645 0.02201 0.01007 -0.0002 1.0000 0.2206
-3.000 -0.3492 0.02078 0.00946 0.0015 1.0000 0.3016
-2.750 -0.3390 0.01892 0.00940 0.0054 1.0000 0.6040
-2.500 -0.2810 0.01948 0.01037 0.0037 1.0000 0.8787
-2.250 -0.1182 0.02043 0.01053 -0.0199 1.0000 0.9886
-2.000 -0.0771 0.02024 0.01008 -0.0234 1.0000 1.0000
-1.750 -0.0707 0.02015 0.00990 -0.0203 1.0000 1.0000
-1.500 -0.0648 0.02010 0.00976 -0.0171 1.0000 1.0000
-1.250 -0.0591 0.02009 0.00967 -0.0139 1.0000 1.0000
-1.000 -0.0535 0.02011 0.00961 -0.0106 1.0000 1.0000
-0.750 -0.0474 0.02017 0.00958 -0.0075 1.0000 1.0000
-0.500 -0.0407 0.02026 0.00959 -0.0045 1.0000 1.0000
-0.250 -0.0332 0.02039 0.00965 -0.0016 1.0000 1.0000
0.000 -0.0040 0.02060 0.00978 -0.0030 0.9942 1.0000
0.250 0.0359 0.02081 0.00992 -0.0063 0.9842 1.0000
0.500 0.0746 0.02101 0.01007 -0.0093 0.9736 1.0000
0.750 0.1130 0.02121 0.01026 -0.0122 0.9633 1.0000
1.000 0.1527 0.02145 0.01049 -0.0152 0.9539 1.0000
1.250 0.1945 0.02169 0.01077 -0.0186 0.9453 1.0000
1.500 0.2376 0.02195 0.01109 -0.0221 0.9368 1.0000
1.750 0.2797 0.02221 0.01144 -0.0254 0.9273 1.0000
2.000 0.3170 0.02248 0.01182 -0.0277 0.9149 1.0000
2.250 0.3327 0.02271 0.01213 -0.0257 0.8940 1.0000
2.500 0.3631 0.02274 0.01228 -0.0263 0.8770 1.0000
2.750 0.4012 0.02268 0.01238 -0.0280 0.8624 1.0000
3.000 0.4446 0.02241 0.01232 -0.0302 0.8465 1.0000
3.250 0.4810 0.02201 0.01211 -0.0307 0.8244 1.0000
3.500 0.5112 0.02163 0.01188 -0.0298 0.7959 1.0000
3.750 0.5403 0.02118 0.01160 -0.0284 0.7637 1.0000
4.000 0.5682 0.02072 0.01123 -0.0266 0.7237 1.0000
4.250 0.5918 0.02044 0.01097 -0.0242 0.6732 1.0000
4.500 0.6125 0.02035 0.01079 -0.0213 0.6072 1.0000
4.750 0.6301 0.02059 0.01065 -0.0180 0.5030 1.0000
5.000 0.6407 0.02152 0.01082 -0.0142 0.3707 1.0000
5.250 0.6478 0.02296 0.01150 -0.0108 0.2627 1.0000
5.500 0.6581 0.02443 0.01246 -0.0083 0.1975 1.0000
5.750 0.6716 0.02576 0.01353 -0.0061 0.1623 1.0000
6.000 0.6870 0.02701 0.01464 -0.0042 0.1405 1.0000
6.250 0.7040 0.02819 0.01578 -0.0026 0.1244 1.0000
6.500 0.7243 0.02935 0.01701 -0.0013 0.1129 1.0000
6.750 0.7466 0.03064 0.01832 -0.0003 0.1053 1.0000
7.000 0.7693 0.03185 0.01963 0.0006 0.0975 1.0000
7.250 0.7927 0.03324 0.02108 0.0013 0.0909 1.0000
7.500 0.8179 0.03484 0.02292 0.0019 0.0862 1.0000
7.750 0.8416 0.03655 0.02474 0.0025 0.0826 1.0000
8.000 0.8638 0.03853 0.02678 0.0032 0.0789 1.0000
8.250 0.8818 0.04047 0.02919 0.0046 0.0749 1.0000
8.500 0.8992 0.04268 0.03172 0.0060 0.0722 1.0000
8.750 0.9147 0.04511 0.03449 0.0074 0.0703 1.0000
9.000 0.9283 0.04758 0.03722 0.0089 0.0686 1.0000
9.250 0.9414 0.05018 0.03992 0.0102 0.0665 1.0000
9.500 0.9432 0.05322 0.04345 0.0128 0.0649 1.0000
9.750 0.9405 0.05646 0.04720 0.0156 0.0634 1.0000
10.000 0.9346 0.05993 0.05107 0.0182 0.0624 1.0000
10.250 0.9247 0.06362 0.05509 0.0208 0.0620 1.0000
10.500 0.9091 0.06735 0.05910 0.0233 0.0618 1.0000
10.750 0.8884 0.07121 0.06316 0.0256 0.0618 1.0000
11.000 0.8649 0.07574 0.06787 0.0263 0.0620 1.0000
11.250 0.8390 0.08123 0.07351 0.0251 0.0623 1.0000
11.500 0.8120 0.08804 0.08041 0.0218 0.0628 1.0000
11.750 0.7867 0.09620 0.08861 0.0168 0.0634 1.0000
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Polar data table (+)
Polar graphs
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