NASA SC(2)-0012 AIRFOIL (sc20012-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA SC(2)-0012 AIRFOIL (sc20012-il) Reynolds number: 100,000 Max Cl/Cd: 33.14 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-sc20012-il-100000.txt Download as CSV file: xf-sc20012-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NASA SC(2)-0012 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.6431 0.09835 0.09303 -0.0031 1.0000 0.1767
-9.500 -0.8935 0.05727 0.05017 -0.0206 1.0000 0.0840
-9.250 -0.8880 0.05278 0.04545 -0.0193 1.0000 0.0835
-9.000 -0.8831 0.04872 0.04102 -0.0175 1.0000 0.0833
-8.750 -0.8769 0.04513 0.03696 -0.0154 1.0000 0.0837
-8.500 -0.8663 0.04162 0.03313 -0.0139 1.0000 0.0856
-8.250 -0.8447 0.03986 0.03148 -0.0136 1.0000 0.0889
-8.000 -0.8285 0.03750 0.02881 -0.0122 1.0000 0.0915
-7.750 -0.8121 0.03521 0.02606 -0.0105 1.0000 0.0945
-7.500 -0.7941 0.03290 0.02343 -0.0093 1.0000 0.0982
-7.250 -0.7718 0.03134 0.02188 -0.0086 1.0000 0.1023
-7.000 -0.7510 0.02995 0.02023 -0.0075 1.0000 0.1074
-6.750 -0.7294 0.02819 0.01829 -0.0065 1.0000 0.1119
-6.500 -0.7074 0.02699 0.01712 -0.0056 1.0000 0.1174
-6.250 -0.6863 0.02597 0.01586 -0.0043 1.0000 0.1229
-6.000 -0.6644 0.02455 0.01458 -0.0034 1.0000 0.1287
-5.750 -0.6444 0.02369 0.01365 -0.0020 1.0000 0.1354
-5.500 -0.6248 0.02262 0.01265 -0.0007 1.0000 0.1418
-5.250 -0.6075 0.02192 0.01197 0.0011 1.0000 0.1500
-5.000 -0.5910 0.02107 0.01125 0.0029 1.0000 0.1583
-4.750 -0.5753 0.02039 0.01062 0.0049 1.0000 0.1692
-4.500 -0.5602 0.01970 0.01007 0.0069 1.0000 0.1835
-4.250 -0.5454 0.01894 0.00954 0.0088 1.0000 0.2036
-4.000 -0.5307 0.01803 0.00896 0.0107 1.0000 0.2436
-3.750 -0.5208 0.01617 0.00827 0.0130 1.0000 0.3929
-3.500 -0.5148 0.01554 0.00906 0.0184 1.0000 0.6725
-3.250 -0.5001 0.01595 0.00951 0.0220 1.0000 0.7357
-3.000 -0.4848 0.01637 0.00994 0.0254 1.0000 0.7740
-2.750 -0.4698 0.01675 0.01030 0.0289 1.0000 0.8051
-2.500 -0.4545 0.01714 0.01067 0.0324 1.0000 0.8301
-2.250 -0.4391 0.01752 0.01101 0.0359 1.0000 0.8541
-2.000 -0.4231 0.01788 0.01134 0.0392 1.0000 0.8788
-1.750 -0.4021 0.01832 0.01173 0.0416 1.0000 0.9027
-1.500 -0.3644 0.01899 0.01231 0.0409 1.0000 0.9246
-1.250 -0.2953 0.02000 0.01322 0.0341 1.0000 0.9458
-1.000 -0.1915 0.02105 0.01412 0.0204 1.0000 0.9653
-0.750 -0.1259 0.02131 0.01432 0.0125 1.0000 0.9774
-0.500 -0.0786 0.02138 0.01436 0.0075 1.0000 0.9863
-0.250 -0.0337 0.02146 0.01441 0.0028 1.0000 0.9946
0.000 0.0000 0.02151 0.01446 0.0000 1.0000 1.0000
0.250 0.0337 0.02145 0.01441 -0.0028 0.9946 1.0000
0.500 0.0785 0.02138 0.01436 -0.0075 0.9863 1.0000
0.750 0.1258 0.02130 0.01432 -0.0125 0.9774 1.0000
1.000 0.1916 0.02104 0.01412 -0.0204 0.9653 1.0000
1.250 0.2952 0.02000 0.01321 -0.0341 0.9458 1.0000
1.500 0.3643 0.01898 0.01231 -0.0409 0.9246 1.0000
1.750 0.4020 0.01832 0.01172 -0.0416 0.9027 1.0000
2.000 0.4230 0.01788 0.01133 -0.0392 0.8789 1.0000
2.250 0.4390 0.01752 0.01101 -0.0359 0.8542 1.0000
2.500 0.4544 0.01714 0.01066 -0.0324 0.8302 1.0000
2.750 0.4697 0.01675 0.01029 -0.0289 0.8052 1.0000
3.000 0.4847 0.01637 0.00994 -0.0254 0.7740 1.0000
3.250 0.5000 0.01595 0.00951 -0.0219 0.7358 1.0000
3.500 0.5147 0.01553 0.00906 -0.0184 0.6728 1.0000
3.750 0.5208 0.01615 0.00827 -0.0130 0.3936 1.0000
4.000 0.5306 0.01803 0.00895 -0.0107 0.2438 1.0000
4.250 0.5453 0.01893 0.00953 -0.0088 0.2036 1.0000
4.500 0.5600 0.01970 0.01006 -0.0068 0.1835 1.0000
4.750 0.5752 0.02038 0.01061 -0.0049 0.1692 1.0000
5.000 0.5908 0.02107 0.01124 -0.0029 0.1583 1.0000
5.250 0.6073 0.02191 0.01196 -0.0011 0.1500 1.0000
5.500 0.6247 0.02262 0.01265 0.0007 0.1418 1.0000
5.750 0.6443 0.02368 0.01365 0.0020 0.1354 1.0000
6.000 0.6643 0.02454 0.01458 0.0034 0.1287 1.0000
6.250 0.6862 0.02597 0.01586 0.0043 0.1230 1.0000
6.500 0.7073 0.02699 0.01712 0.0056 0.1174 1.0000
6.750 0.7293 0.02818 0.01829 0.0065 0.1119 1.0000
7.000 0.7510 0.02994 0.02022 0.0075 0.1074 1.0000
7.250 0.7718 0.03134 0.02188 0.0086 0.1023 1.0000
7.500 0.7941 0.03291 0.02343 0.0093 0.0982 1.0000
7.750 0.8122 0.03520 0.02606 0.0105 0.0945 1.0000
8.000 0.8286 0.03750 0.02881 0.0122 0.0915 1.0000
8.250 0.8448 0.03985 0.03147 0.0136 0.0888 1.0000
8.500 0.8664 0.04163 0.03314 0.0139 0.0856 1.0000
8.750 0.8770 0.04514 0.03696 0.0154 0.0838 1.0000
9.000 0.8832 0.04872 0.04103 0.0175 0.0833 1.0000
9.250 0.8881 0.05279 0.04546 0.0193 0.0835 1.0000
9.500 0.8937 0.05728 0.05018 0.0206 0.0840 1.0000
9.750 0.8170 0.06919 0.06343 0.0252 0.1003 1.0000
10.000 0.8562 0.07120 0.06516 0.0249 0.0972 1.0000
10.250 0.7357 0.10055 0.09528 0.0157 0.1673 1.0000
10.500 0.6322 0.11205 0.10662 -0.0025 0.1594 1.0000
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Polar data table (+)
Polar graphs
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