NASA SC(2)-0010 AIRFOIL (sc20010-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NASA SC(2)-0010 AIRFOIL (sc20010-il) Reynolds number: 50,000 Max Cl/Cd: 22.9 at α=3.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-sc20010-il-50000-n5.txt Download as CSV file: xf-sc20010-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA SC(2)-0010 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
------ -------- --------- --------- -------- -------- --------
-11.000 -0.7184 0.10294 0.09525 -0.0071 1.0000 0.0589
-10.750 -0.7268 0.09599 0.08831 -0.0113 1.0000 0.0586
-10.500 -0.7407 0.08910 0.08139 -0.0157 1.0000 0.0582
-10.250 -0.7588 0.08272 0.07496 -0.0194 1.0000 0.0579
-10.000 -0.7791 0.07724 0.06938 -0.0213 1.0000 0.0576
-9.750 -0.7988 0.07254 0.06454 -0.0213 1.0000 0.0575
-9.500 -0.8136 0.06780 0.05955 -0.0210 1.0000 0.0575
-9.250 -0.8239 0.06319 0.05461 -0.0202 1.0000 0.0579
-9.000 -0.8307 0.05868 0.04961 -0.0190 1.0000 0.0586
-8.750 -0.8263 0.05513 0.04582 -0.0180 1.0000 0.0599
-8.500 -0.8134 0.05270 0.04330 -0.0172 1.0000 0.0617
-8.250 -0.8028 0.04974 0.04003 -0.0162 1.0000 0.0633
-8.000 -0.7911 0.04652 0.03641 -0.0151 1.0000 0.0647
-7.750 -0.7770 0.04344 0.03288 -0.0140 1.0000 0.0668
-7.500 -0.7612 0.04057 0.02935 -0.0127 1.0000 0.0701
-7.250 -0.7411 0.03848 0.02723 -0.0122 1.0000 0.0733
-7.000 -0.7198 0.03649 0.02502 -0.0115 1.0000 0.0768
-6.750 -0.6976 0.03454 0.02263 -0.0106 1.0000 0.0821
-6.500 -0.6744 0.03281 0.02088 -0.0101 1.0000 0.0869
-6.250 -0.6503 0.03132 0.01922 -0.0095 1.0000 0.0933
-6.000 -0.6256 0.02985 0.01766 -0.0090 1.0000 0.1001
-5.750 -0.6005 0.02863 0.01632 -0.0083 1.0000 0.1084
-5.500 -0.5758 0.02743 0.01515 -0.0077 1.0000 0.1170
-5.250 -0.5513 0.02636 0.01399 -0.0070 1.0000 0.1275
-4.750 -0.5061 0.02437 0.01205 -0.0051 1.0000 0.1549
-4.500 -0.4861 0.02332 0.01117 -0.0040 1.0000 0.1746
-4.250 -0.4671 0.02223 0.01029 -0.0028 1.0000 0.2052
-4.000 -0.4504 0.02094 0.00946 -0.0014 1.0000 0.2644
-3.750 -0.4401 0.01937 0.00883 0.0012 1.0000 0.3979
-3.500 -0.4285 0.01871 0.00942 0.0063 1.0000 0.6165
-3.250 -0.4136 0.01878 0.00952 0.0099 1.0000 0.7116
-3.000 -0.3963 0.01894 0.00967 0.0134 1.0000 0.7690
-2.750 -0.3750 0.01924 0.00996 0.0165 1.0000 0.8166
-2.500 -0.3474 0.01962 0.01025 0.0184 1.0000 0.8587
-2.250 -0.2959 0.02025 0.01069 0.0164 1.0000 0.9028
-2.000 -0.2235 0.02063 0.01080 0.0092 1.0000 0.9327
-1.750 -0.1728 0.02055 0.01054 0.0046 1.0000 0.9485
-1.500 -0.1324 0.02037 0.01026 0.0015 1.0000 0.9607
-1.250 -0.0962 0.02020 0.01000 -0.0011 1.0000 0.9719
-1.000 -0.0605 0.02002 0.00978 -0.0036 1.0000 0.9819
-0.750 -0.0251 0.01987 0.00960 -0.0062 1.0000 0.9913
-0.500 0.0066 0.01976 0.00948 -0.0083 1.0000 1.0000
-0.250 0.0042 0.01977 0.00950 -0.0043 1.0000 1.0000
0.000 0.0000 0.01977 0.00951 0.0000 1.0000 1.0000
0.250 -0.0042 0.01976 0.00950 0.0043 1.0000 1.0000
0.500 -0.0066 0.01976 0.00948 0.0083 1.0000 1.0000
0.750 0.0251 0.01986 0.00959 0.0062 0.9913 1.0000
1.000 0.0604 0.02002 0.00978 0.0036 0.9819 1.0000
1.250 0.0962 0.02019 0.01000 0.0011 0.9719 1.0000
1.500 0.1324 0.02037 0.01025 -0.0015 0.9607 1.0000
1.750 0.1728 0.02055 0.01054 -0.0046 0.9485 1.0000
2.000 0.2235 0.02063 0.01080 -0.0092 0.9327 1.0000
2.250 0.2959 0.02025 0.01069 -0.0164 0.9028 1.0000
2.500 0.3473 0.01961 0.01024 -0.0184 0.8587 1.0000
2.750 0.3750 0.01924 0.00995 -0.0165 0.8167 1.0000
3.000 0.3963 0.01894 0.00967 -0.0134 0.7691 1.0000
3.250 0.4135 0.01878 0.00952 -0.0099 0.7116 1.0000
3.500 0.4285 0.01871 0.00942 -0.0063 0.6165 1.0000
3.750 0.4401 0.01937 0.00883 -0.0012 0.3980 1.0000
4.000 0.4504 0.02094 0.00946 0.0014 0.2645 1.0000
4.250 0.4671 0.02223 0.01029 0.0028 0.2052 1.0000
4.500 0.4861 0.02332 0.01117 0.0040 0.1746 1.0000
4.750 0.5061 0.02437 0.01205 0.0051 0.1549 1.0000
5.250 0.5513 0.02636 0.01399 0.0070 0.1275 1.0000
5.500 0.5758 0.02742 0.01515 0.0077 0.1170 1.0000
5.750 0.6005 0.02863 0.01632 0.0083 0.1084 1.0000
6.000 0.6256 0.02985 0.01766 0.0090 0.1001 1.0000
6.250 0.6503 0.03132 0.01922 0.0095 0.0933 1.0000
6.500 0.6744 0.03281 0.02088 0.0101 0.0869 1.0000
6.750 0.6976 0.03454 0.02263 0.0106 0.0821 1.0000
7.000 0.7198 0.03649 0.02502 0.0115 0.0768 1.0000
7.250 0.7412 0.03848 0.02723 0.0122 0.0733 1.0000
7.500 0.7612 0.04057 0.02935 0.0127 0.0701 1.0000
7.750 0.7770 0.04344 0.03288 0.0140 0.0668 1.0000
8.000 0.7911 0.04652 0.03641 0.0151 0.0648 1.0000
8.250 0.8029 0.04974 0.04004 0.0162 0.0633 1.0000
8.500 0.8136 0.05270 0.04330 0.0172 0.0616 1.0000
8.750 0.8264 0.05514 0.04582 0.0180 0.0599 1.0000
9.000 0.8308 0.05868 0.04962 0.0190 0.0586 1.0000
9.250 0.8241 0.06321 0.05462 0.0201 0.0579 1.0000
9.500 0.8137 0.06782 0.05957 0.0209 0.0575 1.0000
9.750 0.7990 0.07257 0.06457 0.0212 0.0574 1.0000
10.000 0.7794 0.07727 0.06942 0.0212 0.0576 1.0000
10.250 0.7592 0.08277 0.07501 0.0193 0.0579 1.0000
10.500 0.7411 0.08916 0.08146 0.0156 0.0582 1.0000
10.750 0.7273 0.09607 0.08838 0.0112 0.0586 1.0000
11.000 0.7190 0.10302 0.09534 0.0069 0.0589 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NASA SC(2)-0010 AIRFOIL (sc20010-il)