NASA SC(2)-0010 AIRFOIL (sc20010-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NASA SC(2)-0010 AIRFOIL (sc20010-il) Reynolds number: 200,000 Max Cl/Cd: 39.04 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-sc20010-il-200000-n5.txt Download as CSV file: xf-sc20010-il-200000-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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.250 -0.9344 0.06621 0.06200 -0.0209 1.0000 0.0206
-12.000 -0.9788 0.05680 0.05223 -0.0256 1.0000 0.0205
-11.750 -1.0042 0.05146 0.04659 -0.0259 1.0000 0.0205
-11.500 -1.0217 0.04761 0.04246 -0.0242 1.0000 0.0206
-11.250 -1.0322 0.04460 0.03917 -0.0216 1.0000 0.0208
-11.000 -1.0337 0.04180 0.03607 -0.0196 1.0000 0.0211
-10.750 -1.0303 0.03918 0.03313 -0.0179 1.0000 0.0216
-10.500 -1.0238 0.03657 0.03017 -0.0162 1.0000 0.0221
-10.250 -1.0140 0.03409 0.02730 -0.0146 1.0000 0.0226
-10.000 -1.0007 0.03186 0.02470 -0.0133 1.0000 0.0231
-9.750 -0.9844 0.03002 0.02264 -0.0122 1.0000 0.0235
-9.500 -0.9658 0.02868 0.02119 -0.0114 1.0000 0.0239
-9.250 -0.9460 0.02752 0.01993 -0.0106 1.0000 0.0245
-9.000 -0.9255 0.02638 0.01867 -0.0099 1.0000 0.0251
-8.750 -0.9044 0.02524 0.01739 -0.0091 1.0000 0.0258
-8.500 -0.8828 0.02410 0.01609 -0.0084 1.0000 0.0267
-8.250 -0.8606 0.02301 0.01483 -0.0077 1.0000 0.0278
-8.000 -0.8385 0.02210 0.01387 -0.0071 1.0000 0.0290
-7.750 -0.8153 0.02143 0.01317 -0.0066 1.0000 0.0304
-7.500 -0.7921 0.02068 0.01234 -0.0060 1.0000 0.0321
-7.250 -0.7687 0.01988 0.01141 -0.0054 1.0000 0.0338
-7.000 -0.7459 0.01911 0.01066 -0.0048 1.0000 0.0356
-6.750 -0.7222 0.01850 0.01002 -0.0042 1.0000 0.0379
-6.500 -0.6982 0.01792 0.00935 -0.0036 1.0000 0.0407
-6.250 -0.6748 0.01734 0.00881 -0.0031 1.0000 0.0439
-6.000 -0.6506 0.01688 0.00831 -0.0026 1.0000 0.0477
-5.750 -0.6272 0.01632 0.00775 -0.0020 1.0000 0.0514
-5.500 -0.6035 0.01591 0.00733 -0.0014 1.0000 0.0558
-5.250 -0.5802 0.01547 0.00688 -0.0007 1.0000 0.0604
-5.000 -0.5577 0.01504 0.00650 0.0001 1.0000 0.0655
-4.750 -0.5354 0.01471 0.00614 0.0010 1.0000 0.0706
-4.500 -0.5107 0.01426 0.00575 0.0013 0.9978 0.0771
-4.250 -0.4766 0.01384 0.00534 -0.0003 0.9903 0.0857
-4.000 -0.4426 0.01344 0.00497 -0.0019 0.9823 0.0973
-3.750 -0.4096 0.01300 0.00463 -0.0033 0.9730 0.1155
-3.500 -0.3775 0.01254 0.00431 -0.0046 0.9634 0.1447
-3.250 -0.3456 0.01200 0.00400 -0.0058 0.9547 0.1957
-3.000 -0.3168 0.01134 0.00371 -0.0065 0.9442 0.2797
-2.750 -0.2899 0.01051 0.00342 -0.0069 0.9344 0.4036
-2.500 -0.2658 0.00969 0.00333 -0.0063 0.9247 0.5617
-2.250 -0.2397 0.00950 0.00343 -0.0056 0.9151 0.6562
-2.000 -0.2115 0.00946 0.00343 -0.0054 0.9069 0.6931
-1.750 -0.1844 0.00943 0.00342 -0.0049 0.8975 0.7204
-1.500 -0.1579 0.00941 0.00345 -0.0042 0.8897 0.7471
-1.250 -0.1316 0.00940 0.00349 -0.0035 0.8812 0.7705
-1.000 -0.1056 0.00940 0.00352 -0.0026 0.8741 0.7906
-0.750 -0.0795 0.00941 0.00357 -0.0019 0.8667 0.8084
-0.500 -0.0531 0.00942 0.00360 -0.0012 0.8600 0.8234
-0.250 -0.0267 0.00943 0.00363 -0.0005 0.8538 0.8361
0.000 0.0000 0.00943 0.00366 0.0000 0.8457 0.8457
0.250 0.0267 0.00943 0.00363 0.0005 0.8360 0.8538
0.500 0.0531 0.00942 0.00360 0.0012 0.8233 0.8600
0.750 0.0795 0.00941 0.00357 0.0019 0.8084 0.8667
1.000 0.1056 0.00940 0.00352 0.0026 0.7907 0.8741
1.250 0.1316 0.00940 0.00349 0.0035 0.7705 0.8811
1.500 0.1579 0.00940 0.00345 0.0042 0.7471 0.8897
1.750 0.1843 0.00943 0.00342 0.0049 0.7205 0.8975
2.000 0.2115 0.00946 0.00343 0.0054 0.6930 0.9069
2.250 0.2397 0.00950 0.00343 0.0056 0.6563 0.9151
2.500 0.2658 0.00969 0.00333 0.0063 0.5624 0.9247
2.750 0.2899 0.01050 0.00342 0.0069 0.4036 0.9344
3.000 0.3168 0.01134 0.00371 0.0065 0.2796 0.9442
3.250 0.3456 0.01200 0.00400 0.0058 0.1957 0.9547
3.500 0.3775 0.01254 0.00431 0.0046 0.1446 0.9635
3.750 0.4097 0.01300 0.00462 0.0033 0.1155 0.9731
4.000 0.4426 0.01343 0.00497 0.0019 0.0973 0.9824
4.250 0.4767 0.01384 0.00534 0.0003 0.0857 0.9904
4.500 0.5108 0.01426 0.00575 -0.0013 0.0771 0.9979
4.750 0.5353 0.01471 0.00613 -0.0010 0.0706 1.0000
5.000 0.5577 0.01504 0.00650 -0.0001 0.0655 1.0000
5.250 0.5802 0.01547 0.00688 0.0007 0.0604 1.0000
5.500 0.6035 0.01591 0.00733 0.0014 0.0558 1.0000
5.750 0.6272 0.01632 0.00775 0.0020 0.0514 1.0000
6.000 0.6506 0.01688 0.00831 0.0026 0.0477 1.0000
6.250 0.6748 0.01734 0.00881 0.0031 0.0439 1.0000
6.500 0.6982 0.01793 0.00935 0.0036 0.0407 1.0000
6.750 0.7222 0.01850 0.01002 0.0042 0.0379 1.0000
7.000 0.7459 0.01911 0.01066 0.0048 0.0356 1.0000
7.250 0.7687 0.01988 0.01141 0.0054 0.0338 1.0000
7.500 0.7921 0.02068 0.01234 0.0060 0.0321 1.0000
7.750 0.8154 0.02143 0.01317 0.0066 0.0303 1.0000
8.000 0.8385 0.02210 0.01386 0.0071 0.0290 1.0000
8.250 0.8607 0.02302 0.01483 0.0077 0.0278 1.0000
8.500 0.8828 0.02410 0.01609 0.0084 0.0267 1.0000
8.750 0.9044 0.02524 0.01739 0.0091 0.0258 1.0000
9.000 0.9256 0.02638 0.01867 0.0099 0.0251 1.0000
9.250 0.9461 0.02752 0.01993 0.0106 0.0245 1.0000
9.500 0.9659 0.02868 0.02119 0.0114 0.0239 1.0000
9.750 0.9845 0.03002 0.02264 0.0122 0.0235 1.0000
10.000 1.0008 0.03187 0.02471 0.0133 0.0231 1.0000
10.250 1.0141 0.03409 0.02730 0.0146 0.0226 1.0000
10.500 1.0240 0.03658 0.03018 0.0162 0.0221 1.0000
10.750 1.0305 0.03918 0.03313 0.0178 0.0216 1.0000
11.000 1.0339 0.04180 0.03607 0.0196 0.0211 1.0000
11.250 1.0324 0.04462 0.03919 0.0215 0.0208 1.0000
11.500 1.0221 0.04762 0.04247 0.0241 0.0206 1.0000
11.750 1.0046 0.05150 0.04663 0.0258 0.0205 1.0000
12.000 0.9793 0.05684 0.05227 0.0254 0.0205 1.0000
12.250 0.9344 0.06636 0.06216 0.0206 0.0206 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NASA SC(2)-0010 AIRFOIL (sc20010-il)